Daniel P. Raymer

Conceptual Research Corporation, Los Angeles, California

Jack Wilson

Arctic Slope Research Corporation, Cleveland, Ohio

H. Douglas Perkins

Glenn Research Center, Cleveland, Ohio

Arthur Rizzi and Mengmeng Zhang

Royal Institute of Technology, Stockholm, Sweden

Alfredo Ramirez Puentes

University of San Buenaventura, Bogota, Columbia

Advanced Technology Subsonic Transport Study

N+3 Technologies and Design Concepts

NASA/TM—2011-217130

November 2011
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question via the Internet to help@

sti.nasa.gov

• Fax your question to the NASA STI Help Desk 

at 443–757–5803

• Telephone the NASA STI Help Desk at

 443–757–5802

• Write to:

 NASA Center for AeroSpace Information (CASI)

 7115 Standard Drive

 Hanover, MD 21076–1320
Daniel P. Raymer

Conceptual Research Corporation, Los Angeles, California

Jack Wilson

Arctic Slope Research Corporation, Cleveland, Ohio

H. Douglas Perkins

Glenn Research Center, Cleveland, Ohio

Arthur Rizzi and Mengmeng Zhang

Royal Institute of Technology, Stockholm, Sweden

Alfredo Ramirez Puentes

University of San Buenaventura, Bogota, Columbia

Advanced Technology Subsonic Transport Study

N+3 Technologies and Design Concepts

NASA/TM—2011-217130

November 2011

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC08VF71P
Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 

Acknowledgments

Funding for this effort was provided by NASA Glenn Research Center. Thanks to Daniel Paxson and Michael Tong of

NASA Glenn Research Center for their technical contributions to this study.
NASA/TM—2011-217130 iii 

Contents 

Abstract ......................................................................................................................................................... 1

1.0 Introduction .......................................................................................................................................... 3

2.0 Nomenclature ....................................................................................................................................... 4

3.0 Background .......................................................................................................................................... 5

4.0 Approach .............................................................................................................................................. 6

5.0 Calibration Basepoint: Boeing 737-800 ............................................................................................... 7

6.0 Parametric Trades and Response Surface .......................................................................................... 10

7.0 Parametric Results Discussion ........................................................................................................... 13

8.0 Advanced Concepts Design ............................................................................................................... 15

9.0 Advanced Concepts Assessment ....................................................................................................... 19

10.0 Preferred Concept—Tailless Airliner ................................................................................................ 21

10.1 Concept Selection and Sizing ................................................................................................... 21

10.2 Enabling Technology: Retracting Canard and Chin Rudder .................................................... 24

10.3 Enabling Technology: Active Aeroelastic Wing ...................................................................... 26

11.0 Tailless Airliner Analysis and Issues ................................................................................................. 27

11.1 Aerodynamics ........................................................................................................................... 27

11.2 Computational Fluid Dynamics (CFD)..................................................................................... 29

11.3 Weights ..................................................................................................................................... 32

11.4 Propulsion ................................................................................................................................. 33

11.5 Sizing ........................................................................................................................................ 35

11.6 Issues and Way-Forward .......................................................................................................... 36

12.0 Summary and Conclusions ................................................................................................................ 38

Appendix—Parametric Sizing Data ............................................................................................................ 41

References ................................................................................................................................................... 42

List of Tables 

TABLE 1.—ADVANCED DESIGN CONCEPTS LIFT ASSESSMENT ................................................ 19

TABLE 2.—ADVANCED DESIGN CONCEPTS SORTING .................................................................. 20

TABLE 3.—NAS2 AERODYNAMIC ANALYSIS .................................................................................. 27

TABLE 4.—UNSIZED NAS2 GROUP WEIGHT STATEMENT ............................................................ 32

TABLE 5.—NAS2 SIZED WEIGHT STATEMENT ................................................................................ 33

TABLE 6.—NAS2 SIZING CALCULATIONS ........................................................................................ 35

TABLE 7.—PARAMETRIC SIZING FOR 4 PARAMETER BOX-BEHNKEN SCHEME 

RESPONSE SURFACE .................................................................................................................... 41

List of Figures

Figure 1.—Preferred concept CAD rendering: Up-and-Away. .................................................................. 3

Figure 2.—Preferred concept CAD rendering: Takeoff. ............................................................................ 5

Figure 3.—Boeing 737 CAD model. .......................................................................................................... 7

Figure 4.—Boeing 737 aerodynamic analysis results. ................................................................................ 8

Figure 5.—Boeing 737 RDS weight estimate. ............................................................................................ 9

Figure 6.—Boeing 737 single variable parametric trades. ........................................................................ 11

Figure 7.—Response Surface equations. .................................................................................................. 12

Figure 8.—Fraction of aircraft design parameter versus baseline. ........................................................... 13

Figure 9.—Advanced design concepts (1). ............................................................................................... 16

Figure 10.—Advanced design concepts (2). ............................................................................................. 18

Figure 11.—Preferred concept Tailless airliner. ....................................................................................... 21
NASA/TM—2011-217130 iv 

Figure 12.—NAS2 vehicle internal arrangement three-view. .................................................................. 22

Figure 13.—NAS2 external arrangement (sized)...................................................................................... 23

Figure 14.—NAS2 showing chin rudder and retracting canard. ............................................................... 24

Figure 15.—NAS2 parasitic drag comparison (unsized NAS2, Ref. to B-737 wing area). ...................... 28

Figure 16.—NAS2 drag savings versus B-737. ........................................................................................ 28

Figure 17.—NAS2 L/D improvement versus B-737. ............................................................................... 29

Figure 18.—NAS2 CFD analysis.............................................................................................................. 30

Figure 19.—NAS2 CFD analysis (beta = 3). .......................................................................................... 30

Figure 20.—Open rotor engine thrust. ...................................................................................................... 34

Figure 21.—Open rotor engine fuel consumption. ................................................................................... 34

Figure 22.—NAS2 on the ramp. ............................................................................................................... 38

Figure 23.—NAS2 three-view. ................................................................................................................. 39
NASA/TM—2011-217130 1 

Advanced Technology Subsonic Transport Study 

N+3 Technologies and Design Concepts 

Daniel P. Raymer 

Conceptual Research Corporation 

Los Angeles, California 91396 

Jack Wilson 

Arctic Slope Research Corporation 

Cleveland, Ohio 44135 

H. Douglas Perkins 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Arthur Rizzi and Mengmeng Zhang 

Royal Institute of Technology 

Stockholm, Sweden 

Alfredo Ramirez Puentes 

University of San Buenaventura 

Bogota, Columbia 

Abstract 

Conceptual Research Corporation, the Science of the Possible, in conjunction with NASA and 

university partners, has completed a two-year study of concepts and technologies for future airliners in the 

180-passenger class. This NASA-funded contract was primarily focused on the ambitious goal of a 

70 percent reduction in fuel consumption versus the market-dominating Boeing 737-800. The study is 

related to the “N+3” contracts awarded in 2008 by NASA’s Aeronautics Research Mission Directorate to 

teams led by Boeing, GE Aviation, MIT, and Northrop Grumman, but with more modest goals and 

funding. 

CRC’s contract featured a predominant emphasis on propulsion and fuel consumption, but since fuel 

consumption depends upon air vehicle design as much as on propulsion technology, the study included 

notional vehicle design, analysis, and parametric studies. Other NASA goals including NOx and noise 

reduction are of long-standing interest but were not highlighted in this study, other than their inclusion in 

the propulsion system provided to CRC by NASA. 

The B-737-800 was used as a benchmark, parametric tool, and design point of departure. It was 

modeled in the RDS-Professional aircraft design software then subjected to extensive parametric 

variations of parasitic drag, drag-due-to-lift, specific fuel consumption, and unsized empty weight. These 

studies indicated that the goal of a 70 percent reduction in fuel consumption could be attained with 

roughly a 30 percent improvement in all four parameters. The results were then fit to a Response Surface 

and coded for ease of use in subsequent trade studies. 

Potential technologies to obtain such savings were identified and discussed. More than 16 advanced 

concept designs were then prepared, attempting to investigate almost every possible emerging concept for 

application to this class airliner. A preliminary assessment of these concepts was done based on their total 

wetted area after design normalization of trimmed maximum lift. 
NASA/TM—2011-217130 2 

This assessment points towards a Tailless Airliner concept which was designed and analyzed in some 

detail. To make it work, a retracting canard and an all-moving chin rudder were employed, along with the 

use of the Active Aeroelastic Wing technology. Results indicate that a 60 percent savings in fuel burn 

may be credibly attained, but this depends upon a lot of technology maturation, concept development, and 

risk reduction. This should be expected—such a dramatic reduction in fuel consumption is a “game 

changer” in the world of commercial aviation. It won’t be easy. 
NASA/TM—2011-217130 3 

1.0 Introduction 

Concepts and technologies to obtain massive reductions in fuel usage in future 180-passenger 

commercial airliners are considered. Funded by a NASA contract related to the ongoing “N+3” efforts of 

NASA and its contractors, Conceptual Research Corporation’s study focused on design concepts and 

technologies needed to obtain a 70 percent reduction in fuel consumption versus current aircraft as 

typified by the Boeing 737. 

Results point to the Open Rotor propulsion system and the Active Aeroelastic Wing as key enabling 

technologies, along with the application of advanced composite, subsystem, flight control, and 

aerodynamic technologies. As far as innovative design concepts, the commonly-considered concepts such 

as Canard, Forward Swept Wing, Blended Wing Body, and Joined Wing do not seem to offer a 

substantial net benefit when the problems of excess wetted area and inadequate trimmed maximum lift are 

included in the assessment. Since a nontrivial reduction in parasitic drag is essential for “N+3” goals, it 

seems that only concepts which actually reduce the total wetted surface area will be winners, leading to 

the tailless approach described below. 

Figure 1.—Preferred concept CAD rendering: Up-and-Away. 
NASA/TM—2011-217130 4 

2.0 Nomenclature 

A Aspect Ratio (span2/reference area, applied to wings and tails) 

AAW Active Aeroelastic Wing 

BWB Blended Wing Body 

CL Wing Lift Coefficient 

CAD Computer-Aided Design 

CFD Computational Fluid Dynamics 

CRC Conceptual Research Corporation 

cg Center of Gravity (mass) 

DATCOM Data Compendium (USAF aerodynamics methodology report) 

FAR Federal Aviation Regulations

L/D Lift-to-Drag Ratio 

M Mach Number 

N+3 Three generations beyond the current commercial transport fleet 

RDS Aerospace vehicle design software (Conceptual Research Corp.) 

SAWE Society of Allied Weight Engineers 

SFC Specific fuel consumption

Swet Aircraft total external wetted surface area 

TRL Technology Readiness Level 

T/W Thrust-to-weight ratio 

W/S Wing loading (weight/area) 

W0 Takeoff Gross Weight 

We Empty Weight 

6-DOF Six Degree of Freedom dynamic simulation 
NASA/TM—2011-217130 5 

3.0 Background 

NASA, since its foundation as NACA in the early days of aviation, has focused research and 

technology development funding on the improvement of efficiency and viability of commercial aircraft. 

In later years it has also spearheaded efforts to reduce aircraft noise and atmospheric pollution, and as 

such can be considered a leading “green” federal agency. The importance of this ongoing research cannot 

be overestimated. Commercial aviation is the largest single positive category in our national balance of 

trade, provides hundreds of thousands of jobs, projects a positive image of a “can-do” America to the 

world at large, and indirectly contributes to our national defense. 

In 2008, NASA awarded $8 million worth of research contracts (Ref. 1) to four industry teams to 

study technologies and concepts for subsonic commercial aircraft in the 2035 time frame. This is 

sometimes referred to as “N+3,” denoting three generations beyond the current transport fleet. 

In the contract press release Juan Alonso, director of NASA’s Fundamental Aeronautics Program, is 

quoted as saying, “The future of air transportation is all about protecting the environment and responding 

to increasing energy costs in a balanced way. We will need airplanes that are quieter and more fuel 

efficient, and cleaner-burning fuels to power them. We are challenging industry to introduce these new 

technologies without impairing the convenience, safety and security of commercial air transportation.” 

Funding was also provided to several NASA Centers. Conceptual Research Corporation was 

contracted by NASA-GRC to consult in matters relating to air vehicle/propulsion design integration and 

to perform independent studies related to the application of N+3 to the 180-passenger class commercial 

transport. This effort is reported herein. 

Figure 2.—Preferred concept CAD rendering: Takeoff. 
NASA/TM—2011-217130 6 

4.0 Approach 

The contract technical phase began with modeling and calibration of the Boeing 737-800 as a pointof-departure baseline. Using the RDS-Professional aircraft design software, CRC modeled the B-737-800 

in sufficient detail to support analysis and alternative design studies. Analysis was done in RDS using 

classical preliminary aircraft analysis methods, mostly those described in this author’s textbook Aircraft 

Design: A Conceptual Approach (Ref. 2). These methods have been extensively calibrated in past studies 

versus actual aircraft including the Lockheed L-1011. For aerodynamics, the methods are primarily based 

on the USAF DATCOM (Ref. 3) plus the Leading Edge Suction induced drag methodology. Weights 

equations are mainly from SAWE1

 papers and statistical equations from Vought Aircraft (Ref. 4). For this 

study, propulsion was based upon data obtained from NASA. 

Once results were obtained and verified with published B-737 numbers, CRC performed parametric 

variations of parasitic drag, drag-due-to-lift, specific fuel consumption, and unsized2

 empty weight. Potential 

technologies to obtain such savings were identified, then a wide variety of advanced concept designs were 

prepared and assessed. A preferred design approach was selected and then refined and analyzed. 

This work was described in two contract-deliverable viewgraph final reports (Refs. 5 and 6) which 

contain more details of the analysis and design efforts. This written report is an attempt to expand and 

present these results to a wider audience in an archival manner, and hopefully influence future concept 

development. 

Be advised that this was a “first-look” study with total funding of only 460 hr, some of which was 

expended in other consulting and engineering support. The budget did not permit extensive analysis such 

as computational fluid dynamics (CFD), structural finite element analysis (FEM), or noise propagation 

estimation. Nor were concepts evaluated in enough detail or by a broad enough cadre of experts to verify 

their feasibility. A follow-on, detailed study is recommended to increase confidence in this study’s 

conclusions and in the viability of the preferred design and technology suite. But, this author is confident 

that the results presented herein are not totally “off base.” 

 1

Society of Allied Weight Engineers 

2

For those not familiar with advanced design methodologies, “sizing” refers to the calculations that determine 

the aircraft takeoff gross weight and fuel weight needed to obtain a specified mission range carrying a certain 

payload weight. This calculation uses numerous design-specific parameters such as the aircraft drag and empty 

weight, leading to a classic “chicken-and-the-egg” problem, namely, you cannot estimate these parameters without a 

design layout, and you cannot do a design layout without the sizing results. To get around this it is common to first 

do an initial design layout based upon a rough guess of the aircraft weight. This is called the “unsized” design, and is 

used to estimate the required parameters for sizing, which is then carried out. From those results, the “unsized” 

layout is scaled and redesigned to a “sized” design which, if all is well, will obtain the desired range carrying the 

desired payload. For details see Raymer (Ref. 2). 
NASA/TM—2011-217130 7 

5.0 Calibration Basepoint: Boeing 737-800 

To impose reality upon such a far-term study it was decide to benchmark all analysis with a realworld, well-known existing commercial aircraft. The Boeing 737 (Ref. 7) was chosen as representative of 

the commercial aircraft fleet. It is the most widely used airliner in the world and over 6,000 have been 

produced. First flown in 1967, it has been updated regularly and incorporates newer engines (CFM-56) 

and in some versions, state-of-the-art winglets for drag reduction. 

A CAD model of the 737 was developed and analyzed to ensure that key range and performance 

values were reasonably matched with pilot handbook data. Aerodynamics analysis was performed, with 

results given in Figures 3 to 5 for a no-winglet configuration. Data for the CFM-56 turbofan engine was 

approximated from NASA estimates, public sources (Ref. 8), and parametric cycle results. Weights were 

estimated using proven empirical weight methods, and adjusted via an input Miscellaneous Empty 

Weight to match the known B-737 empty weight (Fig. 5). This 8,562 lb adjustment is a reasonable value 

for unspecified empty weight items for this class of aircraft. 

A complete FAR commercial transport mission model was developed. Range analysis included 

recalculation of the best cruise altitude and Mach number including thrust and stall limits, as well as 

recalculation of optimal climb schedules, both for this “stock” B-737 calculation and for all parametric 

trade studies and notional concepts to be described. 

Figure 3.—Boeing 737 CAD model. 
NASA/TM—2011-217130 8 

Figure 4.—Boeing 737 aerodynamic analysis results. 
NASA/TM—2011-217130 9 

STRUCTURES GROUP 44896.7 

 Wing 16506.7 

 Horiz. Tail 2010.8 

 Vert. Tail 700.5 

 Fuselage 17940.3 

 Main Lndg Gear 3908.5 

 Nose Lndg Gear 908.7 

 Nacelle Group 2921.1 

PROPULSION GROUP 11271.9 

 Engine(s) 10432.0 

 Engine Controls 42.0 

 Starter 174.9 

 Fuel System 623.0 

EQUIPMENT GROUP 22879.4 

 Flight Controls 3758.5 

 Instruments 202.1 

 Hydraulics 1347.5 

 Electrical 2834.8 

 Avionics 1840.3 

 Furnishings 8147.2 

 Air Conditioning 2074.7 

 Anti Ice 1742.0 

 Handling Gear 52.3 

 APU installed 880.0 

 Misc Empty Weight 8562.0 

 We-Allowance 5.0% 4380.5 

TOTAL WEIGHT EMPTY 91990.5 

USEFUL LOAD GROUP 82209.5 

 Crew 400.0 

 Fuel 38629.2 

 Oil 110.2 

 Payload (max) 13910.0 

 Passengers 29160.0 

TAKEOFF GROSS WEIGHT 174200.0 

Figure 5.—Boeing 737 RDS weight estimate. 

These analytical results were used to calculate range which was used for comparison with the 

published B-737 actual range values. At a maximum takeoff weight of 174,200 lb and an empty weight of 

91,990 lb, takeoff at maximum fuel weight of 46,750 lb (6,875 gal) leaves 35,460 lb for the weight of the 

crew, cargo, passengers, and miscellaneous useful load. This yields a calculated range of 2,774 nm versus 

the B-737 Ops Manual value of 2,820 nm. For the maximum zero fuel weight takeoff (i.e., additional 

cargo and passengers leaving only 35,900 lb for fuel), the calculated range is 2,073 nm versus the 

Ops Manual value of 2,068 nm. 

These close results validate the methods and the model used for subsequent study and analysis. 
NASA/TM—2011-217130 10 

6.0 Parametric Trades and Response Surface 

A main goal of this study was to identify a path for future airliners to obtain a 70 percent fuel savings 

versus the B-737-800. This is not an easy goal, and may prove impossible in that time frame. To better 

understand what it would take to reach it, CRC performed parametric studies on the B-737 baseline 

described above. This sought to determine what contributions must be obtained from savings in vehicle 

weight, aerodynamic drag, and engine specific fuel consumption to add up to such an improvement. In 

such parametric studies, the chosen parameter is “magically” changed to a better value. The aircraft is 

resized to the same airliner mission, resulting in a new sized takeoff gross weight and fuel weight to 

perform the mission carrying the same payload. 

Using this B-737 analytical model, the parasitic drag, induced drag, empty weight, and fuel 

consumption were arbitrarily reduced by 20 and 40 percent and applied in various combinations, sizing 

the aircraft to the benchmark commercial aircraft mission model described above. The maximum fuel 

mission was used as representative of a long-range application. Wing loading (W/S) and thrust-to-weight 

ratio (T/W) were kept constant so that the resized aircraft would attain the required performance values 

for takeoff, climb, and landing. Note that this assumption for T/W may be slightly conservative for some 

cases since the cases that include a drag reduction may be able to attain the required performance with 

somewhat less thrust. 

During this analysis the issue of the proper scaling of aircraft data was addressed. Normally scaling 

effects are negligible since the sized aircraft is usually within 10 to 20 percent of the size of the unsized 

airplane (here the actual B-737). This allows you to assume that the airplane, as it scales up and down in 

size, is essentially “photo-scaled” with all components scaling by the square root of the change in weight 

(so wing loading remains constant). When done this way, the calculated aerodynamic coefficients will be 

unchanged since they are referenced to the wing area. Similarly, it is normally assumed that for small 

changes, the aircraft empty weight will follow historical trend lines as expressed by a straight line on loglog graph paper. When fit to an exponential equation, an exponent of –0.06 is typical. 

Initial runs were made with these normal assumptions, but the results gave airplanes sized far outside 

the 10 to 20 percent variation of normal sizing. Some combinations of reduced drag, weight, and fuel 

consumption gave a sized takeoff gross weight of 50 percent of the actual B-737. Clearly the usual 

assumptions needed reconsideration. 

If the aircraft is scaled down by 50 percent in weight, the photo-scale assumption would have all 

dimensions scale by a factor of 0.707. Obviously the passenger compartment cannot scale by such an 

amount. In fact, the whole fuselage is already about as small as it can be for holding passengers, cargo, 

and the flight crew. Therefore as the aircraft scales down, the total drag coefficient will actually increase, 

since the unchanged fuselage drag value is referenced to the now-smaller wing. 

The same is true for weight—while the weight of wings, tails, tires, and engines may follow historical 

trend lines, the weight of the fuselage will change little as the aircraft gross weight sizes smaller. 

To address the photoscale problem, the drag calculation procedure was modified to allow holding the 

fuselage wetted area constant as the vehicle scales down. This was done as a percent of total wetted area, 

and reasonably simulates the “non-photoscale” sizing effects. 

To develop a credible non-photoscale estimate of empty weight, the B-737 layout was scaled to an 

arbitrary 50 percent reduction in TOGW. The wing, tails, engines, and nacelles were scaled down holding 

T/W, W/S, and tail volume coefficients constant, while the fuselage was left essentially unchanged. 

Weights analysis inputs were revised as appropriate and the resulting empty weight was calculated, and 

then used to define a non-photoscale empty weight log-log trend line equation. This was found to have an 

exponent of –0.31, meaning that when the aircraft’s sized takeoff gross weight is reduced, the empty 

weight does not reduce nearly as much as it would for the photo-scale assumption, and the expected fuel 

improvement is significantly reduced. However, this is more realistic for study purposes, and so it was 

used in all subsequent calculations. 

Using the analysis above and the non-photoscale methods just described, single-variable trade studies 

were conducted and plotted as shown in Figure 6. 
NASA/TM—2011-217130 11 

Figure 6.—Boeing 737 single variable parametric trades. 

Results indicate that a 40 percent improvement in a single parameter reduces the sized aircraft’s fuel 

burn by 49 percent for SFC, by 28 percent for Cdo, by 41 percent for empty weight, and by 21 percent for 

induced drag factor (K). A 25 percent improvement in all four together reduces fuel burn by about 

63 percent. It would take a 30 percent improvement in all four together to reduce fuel burn by our target 

of 70 percent. 

The conclusion to be drawn from this parametric study is that a 70 percent fuel reduction requires 

massive improvements in aircraft characteristics and is unlikely to be accomplished by any one “silver 

bullet.” Significant improvements must be found in all four parameters—parasitic drag, induced drag, 

empty weight, and specific fuel consumption. Of these, the empty weight and engine fuel consumption 

show higher payoff on a percentage basis than aerodynamic improvements but this comparison may be 

meaningless. What is important is where areas for improvements can be found. 

Another conclusion is that the sizing effect leverages the individual improvements but only when they 

are applied to an all-new aircraft design. In other words, savings of this magnitude cannot possibly be 

obtained if improvements are retrofitted to an existing or derivative design. 

To make it easier to estimate the effects of various combinations of parametric improvements to the 

baseline B-737, a Response Surface model was prepared (Ref. 9). First a four-parameter Box-Behnken 

Scheme was devised by NASA’s Jack Wilson, defining specific combinations of variables in increments 

of 25 and 50 percent. These sample cases were analyzed by CRC using the model and methods described 

above, along with single variable trade cases to ensure proper tracking of the obvious variations 

(see Appendix for data). These results were fit to a second degree polynomial Response Surface3

 by 

Wilson using commercial software package SigmaStat, and the resulting equations are provided in 

Figure 7. 

 3

Some third-degree terms are also included. Note that the fuel equation uses the takeoff gross weight response 

surface result as an “input” as well as the parametric variables. 

0

25000

50000

75000

100000

125000

150000

175000

200000

-40 -30 -20 -10 0

% Change

Sized Weight

K

Cdo

C

We

Sized Wo

Sized Wf
NASA/TM—2011-217130 12 

 W0 = 0.551*D + 0.234*K + 0.515*C + 0.0653*D^2 + 0.104*C^2 + 

 0.546*We^2 -0.341*D*K - 0.946*D*C - 0.520*D*We - 

 0.429*K*C -0.463 *C*We + 0.230*K*C*We + 0.917*D*C*We + 

 0.536*D*K*C 

 Wfuel = 0.0718 - 0.154*D + 0.0861*K -0.0722*C - 0.197* W0 - 

 0.0652*K^2 - 0.0522*C^2 + 0.175*D*C + 0.276*D* W0 + 

 0.579*C* W0 + 0.0504*D*K*C + 0.0801*D*K* W0 + 

 0.221*K*C* W0 

Figure 7.—Response Surface equations. 

These equations were programmed by Raymer as routine B737RESP.EXE which quickly determines 

the change in Wo and Wf for input changes in Cdo, K, SFC, and empty weight. All changes are input as 

ratios to the baseline B-737-800 values. Several test cases were conducted to confirm the accuracy of the 

surface fit. Note that these results are specific to the B-737-800 as modeled, using the mission defined 

above. While non-photoscale effects have been incorporated, accuracy will be reduced for solutions far 

from the baseline B-737 size. 
NASA/TM—2011-217130 13 

7.0 Parametric Results Discussion 

Figure 8.—Fraction of aircraft design parameter versus baseline. 

Sample results from the response surface were plotted (Ref. 9) as shown in Figure 8. Note the curve 

labeled “All 4 Simultaneously.” Its intersection with the vertical axis point at 0.3 fraction of fuel burn 

crosses the curve at the 0.7 point on the horizontal axis, namely a 30 percent reduction in each of the 

parametric variables (parasitic drag, induced drag, empty weight, and fuel consumption). In other words, 

we can reach our target of a 70 percent reduction in fuel consumption if we can improve each of these 

design parameters by 30 percent over the baseline B-737 values. 

Finding a way to obtain such savings became the goal of the rest of the study. In the area of Parasitic 

Drag, the obvious places for improvement are reduced wetted area, reduced viscous separation, and 

increased laminar flow. The wetted area can only be reduced by innovative conceptual design. The 

Blended Wing Body approach is often cited for reducing wetted area, but as will be discussed, isn’t 

practical for smaller airliners. To really obtain reduced wetted area in this class of aircraft, components 

need to be eliminated. 

Reduced viscous separation can be attempted by clever shaping, active suction, and the obvious 

approach of thinner wings, tails, and fuselages. All have penalties. 

The baseline B-737 drag buildup using RDS-Professional assumes a typical 10 percent laminar flow. 

If an average of 50 percent laminar flow could magically be attained, the parasitic drag reduces by our 

goal of 30 percent. Increased laminar flow is addressed with smoother fabrication methods, active 

suction, and again, clever shaping. However, attaining 50 percent laminar flow for a real commercial 

airliner is unlikely even in the 2035 time frame. 

The Induced Drag (drag-due-to-lift) depends on the effective span loading, often expressed using 

wing aspect ratio. Quite simply, we need more wing span. However, structural penalties have always 

prevented the use of unusually-large wing spans, along with practical matters such as airport gate and 

taxiway clearances. To address the latter concerns, the Boeing 777 was offered with a folding wingtip 

0

0.2

0.4

0.6

0.8

1

0.5 0.6 0.7 0.8 0.9 1

We

SFC

K

Cd0

All 4 Simultaneously

Fraction of Total Mission Fuel Burn Versus Baseline

Target
NASA/TM—2011-217130 14 

(at great expense to Boeing, but due to the weight penalty, not one airline ordered their planes that way 

including the airline that asked for it!). 

If a 20 percent span increase can be obtained, the airplane sees a 30 percent reduction in induced drag 

since drag goes by span-squared. With current technologies, the wing weight penalty of the greater span 

would be approximately 20 percent, for a total empty weight increase of 4 percent, which must come out 

of fuel. This precludes the simple “more span” approach unless some enabling technology can be found 

which minimizes the weight penalty. 

Winglets4

 and other aerodynamic “tricks” such as spanwise blowing or just a more-optimal spanwise 

lift distribution can possibly help a bit but the main need is simply for span. 

The engine specific fuel consumption also needs a 30 percent improvement to attain our total fuel 

savings goal. SFC depends upon effective disk area, thermodynamic and mechanical efficiency, 

installation losses, propulsion drag, and other factors. Unfortunately, the design cannot tolerate any 

substantial increase in propulsive system weight which rules out certain approaches. 

Of the available approaches to improving propulsive efficiency, the Open Rotor (OR) is probably the 

most promising. Somewhat like an advanced turboprop or turbofan, it increases effective disk area by 

using blades external to the nacelle. Earlier versions called PropFans or Unducted Fans were extensively 

tested and, other than noise issues, proved to be viable and highly efficient. 

Exotic approaches to increase the effective disk area could include remote fans, perhaps electric, 

electrostatic flow acceleration, microactuator blades, or others. Again, potential weight penalties must 

always be considered and will often obliterate any propulsive improvement. 

Along with the Open Rotor, ongoing advances in a variety of propulsion technologies contribute to 

the engine efficiency improvements used in this study. These include an advanced core with an Overall 

Pressure Ratio (OPR) of 70, advanced materials, improved cooling schemes, and greater component 

efficiencies. Should the Open Rotor, for some reason, be unsuitable or less desirable than expected, these 

technologies are fully applicable to an advanced high-bypass turbofan and would yield much of the 

expected benefit. 

The final parameter needing improvement is the vehicle Empty Weight. Unfortunately, many of the 

“good ideas” in the other categories tend to increase empty weight. These must be viewed with extreme 

suspicion, since increases in empty weight have killed more ideas than all other factors combined. 

To actually reduce Empty Weight will take an across-the-board approach. There is no single “long 

pole” item sitting there, waiting to be fixed. The structure of the airplane amounts to only about half the 

total empty weight. Certainly it must be reduced, but such reductions cannot by themselves reach our 

goal. Propulsion, subsystems, and even items such as furnishings must also be reduced. This reduces the 

attractiveness of some of the exotic vehicle designs offering structural weight savings such as the joined 

wing—there simply is not enough weight in the wing structure to “save the day” especially when the 

penalties of those exotic concepts are considered. 

 4

Note that winglets are of marginal benefit for a new aircraft design unless for some reason there is a limit to the 

aircraft’s wingspan. If span is unconstrained, numerous studies have shown that a slightly greater span without 

winglets gives a better aircraft, as proven again during the design studies leading to the B-787. The benefit of 

winglets comes later when an existing aircraft is redesigned to fly at a higher gross weight, and the wing 

spanloading becomes therefore excessive. 
NASA/TM—2011-217130 15 

8.0 Advanced Concepts Design 

Next, the study looked at various exotic vehicle design concepts to assess what improvements if any 

they offer. Using the RDS-Professional aircraft design software, sixteen advanced transport design 

concepts (Figs. 8 and 9) were developed based on the B-737 passenger compartment and payload. These 

were chosen to explore virtually all of the advanced aircraft concept ideas currently under consideration, 

such as forward-swept wings, open rotor, three-surface, blended wing body, joined wing, and more. 

The first concept is a conventional Open Rotor design, similar to Unducted Fan designs done about 

20 years ago. The Open Rotor is perhaps the most logical way to increase propulsive efficiency since it 

increases the total airflow used by the propulsion system with potentially smaller weight and drag 

penalties versus an oversized fan nacelle. Since Open Rotors and similar designs usually cause an 

increase in noise, this first configuration places the exposed rotors all the way at the back of the airplane, 

away from the passenger compartment. This in turn almost forces a T-tail, leading to the unintentional 

“DC-9” appearance. 

The second design places the nacelles on the wings like in a 737. The rotors are in a pusher 

configuration, which might help reduce noise transmission into the passenger compartment. However, 

this location behind the wing can increase the noise generation since the rotor blades have to “cut” 

through the wake of the wing. The alternative tractor location for the propellers is problematic as to 

exhaust location. A much larger portion of total thrust comes from the “jet” portion of the cycle, 

compared to a classic turboprop, so the exhaust flow itself must be efficiently aimed to the rear and kept 

away from other aircraft components like the wing. 

The third design has the Open Rotors at the tail and positioned over an H-tail arrangement to block 

noise transmission to the ground. This is probably even more important than the noise that gets into the 

passenger compartment, if Open Rotors are to find widespread acceptance. 

The Tandem Open Rotor is an original idea for reducing wetted area by placing both engines in a 

single nacelle with counter-rotating rotors. However, this is mechanically complicated, the internal 

ducting would be difficult, and it reduces effective disk area so that some of the propulsion advantage 

may be lost. Still, anything that reduces wetted area is worth looking at. 

Next is a forward swept wing design. FSW was extensively studied in the early 1980’s leading to the 

X-29 flight test program. Advantages include superior wing flow quality, a higher no-flaps maximum lift, 

and better aileron control especially near the stall. Disadvantages include structural divergence, in other 

words, the wing twisting off at a high enough speed. However, this is solvable with tailored composite 

materials as proven in the X-29 program. Unfortunately, FSW gives less flap effectiveness since the flaps 

are highly swept. This forces the wing to be larger than otherwise required. 

This particular design concept adds over-the-wing nacelles located far forward of the wing so that the 

exhaust exit planes are over this wing. This should reduce noise transmitted to the ground. Strut bracing is 

used to reduce the weight penalties of both FSW and the nacelle locations. 

The next two concepts are variants of the Joined Wing (Ref. 10) approach. NASA/Stanford studies in 

the 1980’s indicated a possible 25 percent savings on wing weight due to the bracing effect. 

Disadvantages include aerodynamic interference and the inability to put big flaps on all lifting surfaces. 

The first version shows the “classic” joined wing as envisioned by Julian Wolkovitch. This extends 

the rear bracing wing all the way to the top of the vertical tail, providing maximum triangulation for 

weight savings. However, a separate horizontal tail may be required to provide effective control, so the 

second joined wing concept has such tails. Its rear bracing wings are attached to the nacelles for better 

triangularization and reduced length, hence wetted area. 

Blended Wing-Bodies have been studied by Boeing for many years. The idea is to save wetted area, 

increase lifting efficiency, and reduce weight. The subscale X-48B UAV BWB demonstrator has 

completed initial flight test (Ref. 11) and seems to be validating predicted savings. While the BWB offers 

savings for very large airplanes, for a smaller airplane it is difficult to package the passenger compartment 

without a lot of wasted space. This adds weight and wetted area. 
NASA/TM—2011-217130 16 

Figure 9.—Advanced design concepts (1). 

This design layout assumes 12 abreast in two connected cylindrical compartments, each half the 

length of that in the B-737. These compartments are just as tall as those on a 737 so this BWB does not 

look much like the X-48, and it sees no reduction in wetted area nor improvement in lifting efficiency 

The Oblique wing was invented for supersonic airplanes but offers several benefits for a subsonic 

airliner. It is mechanically simple, has a high critical Mach number, it gives great lift for takeoff and 

landing, and it tends to be light in weight since the wing box is effectively unswept. Of course, there is a 

weight penalty for the pivoting mechanism and the landing gear installation is problematic. 

This design concept places the wing on a short pedestal to minimize interference effects and to avoid 

the possibility of the wing wake entering the inlet duct. 
NASA/TM—2011-217130 17 

The Tandem Wing promises a theoretical 50 percent savings in induced drag, but only if the span is 

kept the same as the equivalent single wing being replaced This is often misunderstood. Drawbacks 

include drag caused by interference between the wings, an increase in total wing weight, problems 

packaging the landing gear, and most important, an inability to put big flaps on the back wing requiring 

the total wing area to be increased substantially. 

The Box Wing was studied in the 60’s but never found an application. It is basically a tandem wing 

with vertical endplates attaching the tips. It promises less induced drag for the same span and supposedly 

has a good transonic area distribution which increases the critical Mach number. Unfortunately, there is 

probably no net structural savings since the geometry doesn’t provide any triangularization. Drawbacks 

include the weight of the wings, the wetted area and weight of the endplates, problems packaging the 

landing gear, and an inability to put big flaps on the back wing. Lockheed studies in the 1960’s also found 

weird flutter modes whose prevention would probably add even more weight. 

The Twin Fuselage concept might offer a slight advantage in wing weight since the total body weight 

is spread out rather than concentrated at the center. However, any benefit would be obliterated by the 

penalties of the extra wetted surface area compared to a single, larger fuselage. In this design the 

fuselages are staggered to mitigate transonic effects, and the tails are mounted only on the more-rearward 

fuselage to maximize moment arm. 

The best way to get a dramatic reduction in drag is to substantially reduce the wetted area. One way 

to do that would be to get rid of the tails, as shown in the next two concepts. In this Tailless design, little 

pop-out canards are provided to trim the pitching moment from the wing flaps, otherwise the airplane 

would be controlled using wing surfaces, wingtip drag rudders, and possibly rudder-like surfaces on the 

engine pylons. Perhaps forebody vortex control could also be used, or at the edge of credulity, cyclic 

control on the open rotor blades. This concept, which became the preferred approach, is described in 

Section 10.0 in greater detail. 

The C-wing is an extension of the winglet, literally a winglet on a winglet. This provides the same 

drag reduction for the winglet that the winglet provides to the wing. As can be seen in the design 

illustration, this last extra surface can also be used as a horizontal tail for pitch control and trim. The 

obvious problem, though, is the structure. These tip-mounted surfaces would cause substantial twisting in 

the wing, especially if used as control surfaces. There will be a severe weight penalty and a great potential 

for flutter. Again, it is worth a look. 

The final option is similar to the Piaggio Avanti, an executive turboprop in production in Italy. This 

uses three surfaces, a wing, an aft tail, and a canard. This provides good pitch control and ease of 

trimming, and also offers a theoretical reduction in trimmed drag. Basically, the downwash from the 

canard cancels out the upwash from the tail resulting in a farfield downwash distribution which can be 

tailored to the optimal elliptical shape. There is also a possible reduction in the wing area since the canard 

is making lift too. It is unclear whether the total aircraft wetted area is reduced by this concept – in this 

study it was not. 

Also, the projected benefits of the three surface design typically apply only in comparison to a 

conventional aircraft design where the tail must create a download for stability and trim. Modern aircraft 

use computerized flight control systems to virtually eliminate this problem. 

(Note: These concepts were previously presented to a non-technical audience in the mass-market 

book Living in the Future (Ref. 12). 
NASA/TM—2011-217130 18 

Figure 10.—Advanced design concepts (2). 
NASA/TM—2011-217130 19 

9.0 Advanced Concepts Assessment 

While funding did not permit an exhaustive analysis and comparison of each of these designs, a 

preliminary study was conducted of two key issues which often dominate any discussion of potential 

improvements: trimmed maximum lift coefficient, and design wetted surface area. These were used for an 

initial ranking. 

First, these designs were all adjusted to approximately the same lifting capability by scaling wing 

areas up or down depending upon estimated trimmed lift capability. Relative lifting abilities were 

estimated based on the total areas of the lifting surfaces, with adjustments based on the likely use of high 

lift devices. A wing surface far to the rear of an aircraft cannot use large flaps—the design can’t be 

trimmed. Such a design would need a greater total lifting area. 

In the absence of a detailed computational aerodynamics study, preliminary estimates of flap 

effectiveness were made based upon historical lift data and approximate balance estimates. It was 

assumed that all lifting surfaces would have the same lift per exposed area without flaps. Then the flap 

effectiveness for each was estimated including trim and nacelle effects. This took into account trim issues, 

with a wing near the center of gravity assumed to have flaps as effective as on a normal airliner, while a 

wing far to the rear would be credited with only 20 percent of the extra lift from flaps5

. The B-737 has 

nacelles on the wing which reduce the trailing edge span available for flaps so concepts with tail-mounted 

engines gain about 10 percent in flap effectiveness. This is shown in Table 1 (file names are the RDS 

design layout files). 

Based on these assessments of relative trimmed maximum lift, each design’s lifting surfaces were 

scaled to give that design the same total lift as the B-737. Tail areas were adjusted in proportion to wing 

area adjustments. Concepts with tail-mounted engines suffered due to the extra wetted area of the nacelle 

pylons, versus the B-737 wing-mounted nacelles which use miniscule stub pylons. 

TABLE 1.—ADVANCED DESIGN CONCEPTS LIFT ASSESSMENT 

Description FileName Sref

Main Wing 

Flap % 

effect.

2nd Wing 

Flap % 

effect.

Relative 

Lifting 

Ability

Baseline 737 B737-800 1255 1.00 0 1.00

OpenRotor Aft Engine NASAWLD1.DSN 1255 1.10 0 1.10

OpenRotor U-tail NASAWL1A.DSN 1255 1.10 0 1.10

Tandem OR U-tail NASAWL1C.DSN 1255 1.10 0 1.10

OR pusher on wing NASAWL1D.DSN 1255 0.95 0 0.95

FSW turbofan NASAWLD2.DSN 1255 0.90 0 1.07

Joined/braced Wing NASAWL3A.DSN 766 1.05 0.3 0.97

Joined Wing NASAWLD3.DSN 766 1.05 0.6 0.88

BWB NASWLD4B.DSN 1255 1.10 0 0.72

Oblique Wing NASAWLD5.DSN 1050 1.20 0 1.00

Tandem Wing NASAWLD6.DSN 630 1.00 0.2 0.88

Box Wing NASAWLD7.DSN 630 1.00 0.2 1.13

Twin Fuselage NASAWLD8.DSN 1255 0.95 0 0.95

Tailless Tandem OR NASAWLD9.DSN 1255 1.10 0 1.10

Tailless OR NASWLD10.DSN 1255 1.10 0 1.10

C-Wing NASWLD11.DSN 1255 1.05 0 1.05

3-Surface NASWLD12.DSN 1255 1.10 0.8 1.21

 

 5

These qualitative assessments are debatable, and a detailed analysis should be made. However, the trim effects 

cannot be ignored. There are canard aircraft that require the wing trailing edge surfaces to be deflected upwards for 

landing, since the wing is so far to the rear. There is definitely a penalty—the question is how much. 
NASA/TM—2011-217130 20 

TABLE 2.—ADVANCED DESIGN CONCEPTS SORTING 

Description FileName Sref

Tailless Tandem OR NASAWLD9.DSN 1255

Tailless OR NASWLD10.DSN 1255

Tandem OR U-tail NASAWL1C.DSN 1255

Joined/braced Wing NASAWL3A.DSN 766

C-Wing NASWLD11.DSN 1255

Baseline 737 B737-800 1255

FSW turbofan NASAWLD2.DSN 1255

OpenRotor U-tail NASAWL1A.DSN 1255

3-Surface NASWLD12.DSN 1255

Tandem Wing NASAWLD6.DSN 630

OR pusher on wing NASAWL1D.DSN 1255

Oblique Wing NASAWLD5.DSN 1050

OpenRotor Aft Engine NASAWLD1.DSN 1255

Joined Wing NASAWLD3.DSN 766

BWB NASWLD4B.DSN 1255

Box Wing NASAWLD7.DSN 630

Twin Fuselage NASAWLD8.DSN 1255

scale factor to 

normalize lift

S-wet after lift 

normalized

S-wet relative 

to baseline

0.909 7422 85.3%

0.909 7843 90.2%

0.909 8548 98.3%

1.029 8565 98.5%

0.952 8585 98.7%

1.000 8696 100.0%

0.931 8700 100.0%

0.909 8780 101.0%

0.826 8797 101.2%

1.141 8814 101.4%

1.053 8897 102.3%

0.996 8910 102.5%

0.909 8920 102.6%

1.140 9188 105.6%

1.385 9424 108.4%

0.883 9734 111.9%

1.052 9796 112.6%

Total wetted areas were recalculated for each design, and then the concepts were sorted by total 

wetted surface area relative to the B-737 baseline (Table 2). Some concepts which appeared beneficial 

based on their initial layout lost their appeal when lift and wetted surface area were taken into account.

The only concepts that significantly reduced the lift-normalized wetted area were those that 

completely eliminated the tails. Most other concepts wound up about the same as the baseline B-737. 

Note that plus or minus 5 percent is probably within the “noise level” for this approximate analysis. 

The best concept was the Tailless design with a single nacelle containing twin tandem open rotor 

engines. This seems to be difficult or impossible to integrate, so it was dropped from consideration in the 

end. The Tailless design with two separate nacelles was next best, showing a potential 10 percent reduction 

in wetted area and hence parasitic drag. It became the baseline for further study as described in this report. 

The worst concepts are those that add a lot of wetted area such as box wing and twin fuselage. The 

Blended Wing Body, developed to reduce wetted area in large aircraft, actually increases it for a smaller 

transport. This occurs because of packaging—it isn’t possible to fit people and cargo into the BWB shape 

without a lot of wasted volume, and that leads to extra wetted area. All the “good” concepts have a 

conventional airliner “cigar” fuselage. 

Another key result is that, unfortunately, none of these advanced concepts appears to be vastly better 

than the baseline Boeing 737, so it appears that an exotic configuration, by itself, is unlikely to “save the 

day.” In fact, it may prove in the end that a traditional Boeing design, looking like the 737, 757, and 767, 

is about as good as you can get. That means that tomorrow’s 737 replacement, like the B-787, may look 

fairly normal from afar while incorporating the latest advanced technologies “under the skin.” 

These comparisons do not yet include the weight effects. Qualitatively, weight tends to be related to 

total wetted surface area so the four concepts with the greatest wetted areas are likely to be heavier as 

well as “draggier.” This is another benefit of getting rid of the tails—missing tails weigh nothing. Of 

course, the unknown mechanisms that must be used to provide control and trim will weigh something. 

Only a detailed study can sort things out. 

Designs with high aspect ratio wings are heavier. This is especially a problem for the tandem wing 

where chord length was halved to maintain span with a similar total wing area. A more detailed study 

may find a more-optimal approach than this theoretical assumption, but it is doubtful that the tandem 

wing will ever look good simply because the back wing cannot use large flaps. 

The Oblique Wing should show a substantial weight savings in the wing itself, since it has zero 

effective structural sweep. The pivot mechanism and landing gear installation will add back some weight. 
NASA/TM—2011-217130 21 

10.0 Preferred Concept—Tailless Airliner 

10.1 Concept Selection and Sizing 

With everything considered, the advanced design concept with the best potential for obtaining 

meaningful improvements seems to be the Tailless approach. It alone provides a substantial reduction in 

wetted surface area once the wing size is adjusted to provide equivalent lift. Reduced wetted surface area 

leads directly to reductions in both parasitic drag and weight—if this obviously-risky design approach can 

be made to work. The Open Rotor propulsion system, while not essential to the Tailless design approach, 

seems to offer the greatest payoff in terms of engine specific fuel consumption and is included in this 

baseline. 

Of the two Tailless designs shown in Figure 10, the better in terms of wetted area is the one with both 

of its Open Rotor propulsion systems integrated into a common nacelle. Such an approach is not unheard 

of—several large propeller-powered aircraft like the new Dornier Seastar CD2 have used twin engines in 

a single nacelle, back-to-back with a tractor propeller at one end and a pusher propeller at the other. 

Alternatively, the never-produced LearFan flew well with twin, side-by-side turboshaft engines driving a 

single pusher propeller. 

It is possible that one of these approaches could result in a workable tandem nacelle Open Rotor 

approach. However, the greater core massflow of the Open Rotor system would make it difficult to have 

the exhaust from the front engine get safely past the second engine and rear-mounted rotor, so both 

engines were assumed to be at the rear. This still leaves a substantial packaging problem, with twin 

tandem core turbojets, shafting to get mechanical power from each core to its respective Open Rotor, 

perhaps gearboxes, and the ducting associated with intake air and core exhaust flow. In the concept layout 

shown above this was roughly “eyeballed” but subsequent review by CRC and NASA indicates that such 

packaging is closer to “impossible” than merely “difficult.” 

Thus, the twin-nacelle Tailless Open Rotor design was selected as the most-promising approach with a 

reasonable chance of success. A detailed configuration design layout of this preferred concept was made 

using RDS-Professional. First an “unsized” layout was done (NAS1, not shown), assuming the same takeoff 

gross weight as the B-737-800. The fuselage geometry was a virtual copy in terms of length and diameter, 

with slight reshaping to accommodate the retracting canard. The engines were analytically scaled to provide 

equivalent thrust. Wing area of the unsized design matches that of the B-737, as do tire sizes. 

Figure 11.—Preferred concept Tailless airliner. 
NASA/TM—2011-217130 22 

The wing aspect ratio is substantially higher than the B-737. This offers a major improvement to the 

drag-due-to-lift (induced drag) but is normally not permitted due to the weight penalty. As discussed in 

Section 10.3, the Active Aeroelastic Wing (AAW) technology permits a higher aspect ratio with minimal 

weight penalty, and is potentially a key enabler for N+3 aircraft. Note that the total wing span of 115 ft, 

despite the higher aspect ratio, is only 2.6 ft (2 percent) greater than the B-737 once sizing effects are 

taken into account that follows. 

This unsized layout was then analyzed as to aerodynamics and weights. Along with NASA-provided 

propulsion data, these results were used to resize the airplane, yielding a takeoff gross weight (Wo) of 

122,592 lb and a fuel weight (Wf) of 17,203 lb. This sizing should allow the airplane to carry the same 

payload over the same FAR mission range as the B-737-800. The baseline design (NAS2) was redrawn 

accordingly, as shown in Figures 12 and 13. The fuselage stayed about the same, but the wings, landing 

gear, engines, and canards were reduced in size since the aircraft weight was reduced. A description of the 

features of this baseline design follows. 

The NAS2 preferred concept is a normal airliner in most regards, with a conventional airliner 

fuselage, passenger compartment, cargo bay, cockpit, landing gear, and aft-fuselage-mounted engines. In 

some ways it looks like the old DC-9. The big difference, of course, is the absence of both vertical and 

horizontal tails. 

Its fuselage would be fabricated from next-generation advanced composites. The passenger and cargo 

arrangements are similar to the B-737, including doors, galleys, toilets, and all the mundane details of 

airliner design. Landing gear is conventional, but presumably lighter in weight using composites or 

advanced metallics. 

The two-man crew station is conventional in arrangement but would incorporate the latest in display 

and control technology. By that time frame a one-man crew might be used, with sufficient automation to 

land the plane should the pilot become incapacitated. Even an unpiloted airliner may occur by this time 

frame but in all the vehicle analysis that follows, a two-man crew was assumed. 

Figure 12.—NAS2 vehicle internal arrangement three-view. 
NASA/TM—2011-217130 23 

Figure 13.—NAS2 external arrangement (sized). 
NASA/TM—2011-217130 24 

10.2 Enabling Technology: Retracting Canard and Chin Rudder 

The obvious difference in this design compared to current airliners is the aft tails. There are not any. 

This saves a lot of weight and drag (wetted surface area) but so far, all airplanes of this configuration have 

needed tails. 

Eliminating tails provides much of the benefit of the flying wing. The Blended Wing Body discussed 

above is a direct result of attempts to provide these flying wing benefits in a design with sufficient room 

for passengers. Unfortunately, packaging considerations make the BWB a poor choice for smaller 

airliners. There is just too much wasted space. This tailless design can be seen as an attempt to get the 

same benefit—a reduction in wetted area and weight—in an airliner with tighter internal packaging. 

Tails are needed for three things—stability, control, and trim. With a modern flight control system, 

stability is unneeded as long as sufficient control is attained. This design philosophy has been in use since 

the F-16 for military aircraft. The N+3 timeframe will be 50 years after the F-16 entered production. It is 

time for that technology, like so many military technologies before it, to be accepted for airliners. So 

forget stability—we just need control. 

However, we must also make sure that trimming the aircraft under widely-variable loading conditions 

does not cause excessive trim drag. Airliners see wider ranges of payload cg location than do fighters and 

bombers. To provide trim, the nacelle pylons have movable “ruddervators” at the trailing edge. These are 

also used for additional pitch and yaw control when needed. 

In addition, the wing itself will provide pitch and yaw control as well as the traditional roll control. This 

is done using the Active Aeroelastic Wing (AAW) technology, described in a forth coming section. 

A retracting canard horizontal tail surface is provided, mostly to counter the huge nose-down pitching 

moments created by the large wing flaps. On a normal, aft-tail configuration, the horizontal tail deflects to 

an extreme angle when flaps are extended, creating a download just when additional lift is desired. This 

retracting canard was sized by these flaps-down moments, with a 15 percent lift excess for control 

purposes. Extending the retractable canard also adds lifting surface for takeoff and landing, allowing the 

actual wing to be smaller for reduced drag during cruise. 

This canard mechanism, when extended, is similar to that of the Boeing Sonic Cruiser where a lot of 

technology development work was put into exactly such a system. This design is similar, but adds 

retraction to the problem. This would be similar to the wing pivot on the B-1 or F-14. Luckily, “N+3” 

means that there is a lot of time to develop practical mechanisms. 

Figure 14.—NAS2 showing chin rudder and retracting canard. 
NASA/TM—2011-217130 25 

The canard also provides pitch control and trim during takeoff and landing. When the canard is 

retracted, these are provided by a combination of wing moments (like the B-2) and nacelle pylon 

ruddervators. 

Note that this design is not simply a new form of the rarely-used lifting canard configuration. In our 

design the canard retracts completely for efficient cruising flight. In the Sonic Cruiser and recent studies 

by a major contractor, the design is a LIFTING canard. This acts much like a tandem wing and, like a 

tandem wing, is less efficient for lifting than a wing by itself. For stability reasons the canard of such a 

design is usually carrying more than its proportional share of the weight, leading to inefficiencies. This 

problem with lifting canards has been known since the 1930’s by airplane designers, and is why that type 

is rarely seen. The NAS2 design reported herein is a tailless design during cruising flight, with its pop-out 

canard only used when wing flaps are deployed. 

A computational aerodynamic study with detailed 6-DOF dynamic simulation is required to properly 

assess canard sizing and elevator actuation requirements. Hand calculations were used to initially size the 

required surfaces. 

For yaw control, the preliminary layout assumed a to-be-determined solution. After considering 

wingtip drag devises, vectored thrust, and forebody vortex control, an all-moving chin rudder was 

designed for the actual NAS2 layout. Since it is farther from the aircraft center of gravity than an aft 

vertical tail, it can provide the same yawing control with a smaller area. Being all-moving, it provides 

more control authority and can be made even smaller. 

While the chin rudder could also be retracting, it is probably not worth the trouble and was left fixed 

in this design. It was sized to provide adequate yaw authority for an engine-out case, with a 10 percent 

force margin. 

Nacelle pylon ruddervators, along with wingtip drag control Section 11.1, provide an additional 

margin for extreme yaw upsets. This too needs a full analysis and dynamic simulation. 

Both the horizontal canard and the vertical chin rudder make the aircraft unstable, requiring highspeed actuation and a sophisticated computer system. While ambitious for a commercial airliner, such 

technology has been flight tested in the pitch axis numerous times, most notably on the X-31. Applying 

the same technology to the yaw axis should pose no insurmountable problem in the N+3 timeframe. 

The X-31 program also demonstrated the technologies for flight without a vertical tail, but in a manner 

not quite identical to that proposed here. In the “Quasi-Tailless Flight Demonstration” program (Ref. 13), 

the X-31’s computerized flight control system was modified to have the rudder and ailerons deliberately 

negate the stabilizing effects of the vertical tail. Then the nozzle vectoring system was used to stabilize the 

aircraft in yaw. With the tail fully removed, aerodynamic lateral stability would go from 0.0040 to –0.0018 

at Mach 1.2 which is beyond the yaw authority of the nozzle vectoring so a “partial tail removal” scheme 

was implemented. Flight test reached 70 percent effective tail removal, a good demonstration but not quite 

full tailless flight. However, the X-31 was not originally designed for tailless flight and has no suitable 

controls other than the nozzle, which was not designed with tailless flight in mind. 

More recently the McDonnell Douglas/Boeing X-36, having no vertical tail, did fully demonstrate 

tailless flight (Ref. 14). X-36 was naturally unstable in the yaw axis and used an advanced thrust 

vectoring nozzle for directional control. In 1997 this unmanned technology demonstrator made 31 flights 

and is said to have flown quite well. 

The relevant lesson-learned from these flight test programs is that such tailless flight using artificial 

stability is indeed possible, but it is necessary to have sufficient control authority. In the NAS2 design this 

is provided by the chin rudder with augmentation as needed from the pylon ruddervators and the 

application of wingtip drag via the Active Aeroelastic Wing. An important objective of future studies will 

be to perform CFD and 6-DOF studies in enough detail to confidently size the chin rudder to meet all 

contingencies in a manner consistent with flight safety and relevant FAR’s. This is further discussed in 

the CFD Section 11.2. 
NASA/TM—2011-217130 26 

10.3 Enabling Technology: Active Aeroelastic Wing 

One of the most exciting technologies that could be applied to N+3 airliners is the Active Aeroelastic 

Wing (AAW) (Ref. 15). Developed by Rockwell International (now Boeing) during the Advanced 

Tactical Fighter (F-22) program and flight tested on the X-53, AAW can provide 3-axis flight control, 

modulate drag, alleviate gust loadings, reduce wing weight, and permit higher aspect ratio wings. 

Originally developed to solve problems with excessive flexibility in Rockwell’s Advanced Tactical 

Fighter design, the AAW was first conceived as simply a computerized implementation of a part of every 

B-47 pilot’s training. As described in Living in the Future (Ref. 12); 

“The B-47 was the first high-speed jet with long, thin, skinny swept wings. It had such bad roll 

reversal that Boeing had to add wing spoilers. Above 450 kts the plane is rolled using spoilers to 

kill some lift on one side. But sometimes the spoilers broke. So the pilots trained to fly the plane 

backwards. They’d take it up to about 500 kts, pull the spoiler circuit breaker, and practice moving 

the stick to the right to roll the big bomber to the left.” 

In the AAW, the wing is deliberately built with reduced torsional stiffness, while of normal strength in 

all other directions. Normally, torsional stiffness requirements substantially increase the weight of an 

airliner’s wings and are frequently the most-critical loads on the outer third or more of the wing. So weight 

is saved, right at the start. 

Next, high-speed actuators are used on both leading edge and trailing edge control surfaces. Rather 

than traditional airliner slats, the leading edge movable surfaces are designed like those for fighter wings, 

pivoting flaps that can be deflected at low and high speeds. They can deflect up as well as down. 

Finally, a computerized control system is implemented to deflect those surfaces properly. For roll 

control, the trailing edge surfaces deflect like normal ailerons at low speeds. At higher speeds they deflect 

in the opposite direction. If the trailing edge control surface is deflected downwards it sees increased lift, 

which lifts the back of the wing box and twists the wing leading edge down. This twisting reduces the lift. 

At higher speeds, enough twist is obtained to overpower normal aileron effect and the net force is 

downward. Like the B-47, the airplane rolls in the opposite direction. 

Similarly, downward deflection of the leading edge flaps produces a download at the front of the 

wing box, twisting the wing leading edge down. 

This wing twisting ability can also be used to twist both wings leading edge down, reducing lift out at 

the tips. This can be used for gust alleviation, for optimizing spanwise lift distribution for drag and structural 

relief, and for pitch control much like the elevons of a flying wing. The high-speed gust alleviation will 

reduce wing structural loads, reducing weight even more and allowing a higher aspect ratio wing. 

AAW even allows the creation of additional wingtip drag for yaw control and to act as speed brakes. 

This is done by deflecting the leading and trailing edge surfaces so that they “fight” each other, each 

trying to twist the wing in the opposite direction. 

The Active Aeroelastic Wing is at TRL 6 or 7 for application to commercial airliners. It was flight 

tested on a modified F/A-18 Hornet as a joint project of the Air Force Research Laboratory, Boeing 

Phantom Works and NASA Dryden. The F-18, renamed X-53, was modified to make its wings more 

flexible. It flew in 2002 and made over 50 flights, half of which were supersonic (Ref. 16). 

Given that most commercial airliners must lock out the outboard ailerons in high-speed flight to 

prevent roll reversal, the fruitful application of AAW seems assured. 
NASA/TM—2011-217130 27 

11.0 Tailless Airliner Analysis and Issues 

11.1 Aerodynamics 

While not visible on the drawing, a number of advanced technology aerodynamic improvements were 

assumed for the NAS2 design. An increase in laminar flow from a current 10 percent to an overall 

25 percent assumes the use of CFD for design shaping, smoother skins from composite material, better 

fabrication quality, and possibly active suction in key areas. The flat-glass windshield drag penalty of the 

B-737 was eliminated, as is the case for the B-787. 

In classical airliner aero analysis there is a substantial penalty from the bottom rear of the fuselage, 

where the contour suddenly angles up towards the tail. This “upsweep angle” as used in a common drag 

equation was reduced from the actual 8 to 6 representing improvements in bottom flow separation via 

CFD, optimal design, and possibly local suction. 

The induced drag receives an obvious credit for the high aspect ratio wing, permitted by the use of 

AAW technology. 

Of course, the greatest improvement in aerodynamics comes from the reduced wetted surface area 

brought about by the tailless design approach. 

Aerodynamic analysis for the NAS2 design was done using the classical methods in the RDSProfessional program (Ref. 17). As described above, this analysis was calibrated to the B-737-800 and 

gave good results. Subsonic parasite drag was estimated by the component buildup method. Key 

aerodynamic assumptions and sample parasitic drag results are listed in Table 3. 

 Smooth Paint Equivalent Roughness 

 25 percent Laminar Flow (averaged over whole surface) 

 4 percent Leakage and Protuberance Drag applied to total wetted area drag 

TABLE 3.—NAS2 AERODYNAMIC ANALYSIS 

Altitude = 30000. ft Mach = 0.80 Velocity = 471.5 (kts) 

 # R# 10^6 Cf FF Q Adj S-wet Cdo 

WING 1 19.433 19.921 1.553 1.000 1.000 1954.2 71.2 

FUSELAGE 1 285.230 13.265 1.092 1.000 1.000 4350.9 95.5 

 Includes Upsweep drag= 20.4 

NACELLE 2 32.419 18.322 1.076 1.000 1.000 190.0 8.8 

PYLONS 1 23.590 19.292 1.479 1.000 1.000 171.3 5.8 

 Drag rise or Wave Drag coeff Cdw 13.7 

 TOTAL PARASITE DRAG COEFFICIENT Cdo 195.0 

-Drag is referenced to the theoretical reference wing area: Sref=883 sqft 

-Drag is expressed in "counts" - multiply by 0.0001 to get coefficient 

 
NASA/TM—2011-217130 28 

The following drag results are from the unsized NAS2 design, with the same wing as on the B-737. 

This illustrates the total 33 percent drag savings obtained from these combined improvements. The 

individual contributions are detailed in Figure 16. 

Figure 15.—NAS2 parasitic drag comparison (unsized NAS2, Ref. to B-737 wing area). 

Figure 16.—NAS2 drag savings versus B-737. 

 
NASA/TM—2011-217130 29 

Figure 17.—NAS2 L/D improvement versus B-737. 

The drag-due-to-lift was calculated by the leading-edge suction method using a calculated lift curve 

slope based upon DATCOM methods. A leading-edge suction schedule typical for commercial transports 

was employed, based on the design lift coefficient. 

The ultimate drag measure of merit is the aircraft’s total lift-to-drag ratio (L/D), which includes 

parasitic drag and the drag-due-to-lift. This term alone enters into the classic Breguet Range Equation, 

and the resulting range of an aircraft is directly proportional to L/D. The L/D of the unsized NAS2 is 

shown in Figure 17, showing a 52 percent improvement over the baseline B-737 with the same size wing 

for comparison sake. 

With the limited budget of this project, detailed stability calculations were not permitted. A classical 

pitch stability method estimated a Static Margin of –4.73 at Mach 0.6 (canard retracted), which is a 

reasonable goal for an advanced technology, computer-controlled aircraft. 

Far more work is required in the area of stability and control. A believable answer will require CFD 

coefficient analysis and a 6-DOF aeroelastic analysis with active flight controls in the loop. Based on 

experiences during the design of Rockwell’s highly-flexible Advanced Tactical Fighter concept, this 

author expects ultimate success in the area of control system design including the AAW system. Of 

course, obtaining FAA certification for such a design will be yet another challenge, but such challenges 

must be met if our fuel savings goals are to be attained. 

11.2 Computational Fluid Dynamics (CFD) 

To permit computational analysis not budgeted under contract funding, an arrangement was made 

with Sweden’s Royal Institute of Technology (KTH) wherein the external geometry of the NAS2 was 

provided as a sample case for an ongoing cooperative effort. This effort is aimed at creating a rapid 

interface between Conceptual Research Corporation’s RDS-Professional Design Layout Module and the 

CFD model interface program (Ref. 18) being developed at KTH under the direction of Dr. Arthur Rizzi. 

After loading the geometry and converting it via the interface program, KTH graduate student 

Mengmeng Zhang performed Euler analysis using the Swedish EDGE unstructured CFD solver. 

Additional analysis was done using the TORNADO vortex lattice method (VLM), especially for 

analyzing control effectiveness. These results should be considered preliminary since the goal of the KTH 

researchers was tool development (Ref. 19) and not final aerodynamic calculations. Nevertheless, the 

results are of considerable interest to this project. 
NASA/TM—2011-217130 30 

One calibration point can be extracted from the data which allows comparing the results from the 

classical analysis in RDS to the Euler result. The difference in CL from 0 to 1 at Mach 0.8 in the Euler 

analysis gives a CL-alpha of 0.155 per degree. The RDS classical analysis, mostly based on DATCOM, 

gives CL-alpha = 9.18 per radian, or 0.160 per degree. 

Pressure profiles from the Euler runs are shown in Figures 18 and 19, at Mach 0.8 and 1 angle of 

attack. Nothing unusual is seen, although better streamlining of the cockpit area appears in order. A closer 

investigation showed that the RDS geometry in this area is a bit rough, having originally been based on 

the B-737 cockpit like all of the fuselage. In a subsequent study it should be smoothed out, like on the 

B-787 (which itself is based on the deHaviland Comet, in a full circle of jet aircraft history). 

Figure 18.—NAS2 CFD analysis. 

Figure 19.—NAS2 CFD analysis (beta = 3). 

 
NASA/TM—2011-217130 31 

Given the absence of a traditional vertical tail, the most critical calculation for this exotic design is the 

yaw control via chin rudder. It is assumed that stability is provided artificially, but to make that happen, 

adequate control must be assured. To improve effectiveness, the chin rudder is all-moving and in an ideal 

location for control effectiveness, being both at the front of the aircraft where good flow quality is 

assured, and being at the end of a long moment arm relative to the center of gravity. 

It must be assumed that the chin rudder actuator will be totally fail-safe and will operate at a high 

enough pivot rate and frequency that its time lag will be negligible. This is not a problem for modern flight 

control actuators, being just a matter of cost and weight. The X-31 canard is a good example, and that flew 

over 20 years ago. Of course, one of the X-31’s crashed due to a flight control system failure, but that 

involved a non-redundant, unheated pitot tube which was inadvertently taken into icing conditions. 

When developing the NAS2 design layout, the chin rudder was sized based on countering the yawing 

moments produced during an engine-out event. These include the unbalanced thrust of the remaining 

engine, and the drag of the inoperative engine and its open rotor. These drags were approximated in the 

absence of manufacturer’s data, but should not be too far off. 

However, engine-out control is not the only concern for the design of the chin rudder. As an unstable 

system, the control effector must provide a suffient restoring moment from a sideslip condition. The 

control gain is also important. This is the number of degrees of deflection required to counter a single 

degree of system offset, in this case being a yaw or sideslip. If it takes, say, 5 of rudder deflection to 

create a restoring moment in the event of a 1 yaw upset, the gain is probably too high. The problem is, 

what happens if there is a huge lateral gust or other upset that causes a 10 yaw offset? It would not be 

practical to expect a 50 chin rudder deflection to correct the yaw. The chin rudder probably stalls at 25

deflection or less. 

A conservative rule of thumb for aft control surfaces is to limit the gain to one. A 10 sideslip offset 

would require only a 10 deflection of a traditional rudder. Note that for a traditional aft rudder, attached 

to the aircraft, that 10 of sideslip already provides 10 of rudder deflection relative to the freestream. A 

gain of one is, in effect, a total gain of two. This is the conservative rule of thumb for gain, re-expressed 

in terms of the freestream. 

For an all-moving canard surface (vertical or horizontal), the computer can easily actuate the surface 

to “weathervane.” In other words, if the aircraft nose deflects 10 to the left, the computer automatically 

pivots the chin rudder 10 to the right, so that it remains aligned with the freestream. If it is then pivoted 

an additional 10 to the right for a total of 20 relative to the body axis, there is a mechanical gain of two 

that is, in effect, a gain of only one relative to the free stream. 

While of a preliminary nature, the Euler code analysis indicates that the yawing moment at Mach 0.8 

and 3 of sideslip is about 0.00297. A vortex lattice analysis of the chin rudder gives its yawing moment 

(Cn) per degree as about –0.0005, so about 6 relative to the freestream provides the required moment. 

This is a gain of two relative to the freestream, identical to the “rule of thumb” discussed above. 

Obviously this result is preliminary and is subject to revision as detailed analysis becomes available. 

Still, it is encouraging that an Euler analysis seems to support the basic notion of an unstable tailless 

commercial transport that uses an all-moving chin rudder to provide both yaw control and artificial yaw 

stability. This creative definition of gain is debatable and only a detailed analysis and simulation will 

ultimately answer the question, “is this feasible?” 

These results conservatively assume that the chin rudder is being used alone for yaw control. As 

described above, it would be augmented with differential drag from the Active Aeroelastic Wing, and with 

the ruddervators on the back of the nacelle pylons. These will reduce the gains required for the chin rudder. 

One additional effect that awaits detailed analysis—a conventional aft-rudder suffers from aeroelastic 

effects. The fuselage bending allows deflection of the fixed vertical tail in a direction which reduces net 

yaw control. With a canard rudder the opposite effect is seen. Fuselage bending will actually increase 

control effectiveness as the now-bent fuselage generates its own yawing moments. This effect, probably 

minor, was ignored for this study. 
NASA/TM—2011-217130 32 

11.3 Weights 

The empty weight buildup for the NAS2 was estimated in RDS-Professional using classical methods, 

mostly the Vought Aircraft semi-physics-based empirical transport weight equations (Ref. 4). For 

calibration purposes, these time-honored methods were applied to the B-737-800 as described above, and 

provided what seem to be reasonable answers. Unfortunately, direct comparison was not possible since 

the actual empty weight statement of the real airplane is considered company proprietary even after all 

these years. 

For the NAS2, the design geometry was used along with various parameter improvements to the 

weight model representing advanced technologies. In the absence of a detailed FEM analysis, the 

reduction in wing weight due to the use of the Active Aeroelastic Wing was approximated by analogy. It 

was assumed that the AAW technology would minimize the need for torsional stiffness and would also be 

used to provide maneuver load control, gust alleviation, and flutter suppression, thus countering the 

weight penalty normally associated with increased aspect ratio. The B-737 wing aspect ratio was 

parametrically increased by 50 percent and its weight recalculated. An 18 percent weight reduction was 

found to negate the weight penalty, and was then applied to the NAS2 wing. 

Weight reduction factors were applied to various aircraft structures to account for the use of nextgeneration advanced composite materials, namely 15 percent for the wing, 20 percent for the canard, 

18 percent for the fuselage, and 20 percent for the pylon. These are slight improvements over the values 

that have been widely used for current technology composites. Systems, furnishings, and avionics weights 

were reduced 25 percent versus B-737 technology levels. 

The additional weight penalty for the retracting canard was estimated by analogy to variable-sweep 

wings, which normally suffer a 19 percent penalty. The chin rudder calculated weight was increased by 

20 percent to account for mechanisms and stress concentrations associated with an all-moving surface. 

A 5 percent empty weight margin was then applied on top of the total estimated empty weight. The 

resulting Group Weight Statement for the unsized aircraft is Table 4. Recall that this is the initial layout 

using the original B-737 wing area, thrust, tire size, and such, and shows the design and technology 

savings for direct comparison. 

TABLE 4.—UNSIZED NAS2 GROUP WEIGHT STATEMENT 

Weight Loc Moment Weight Loc Moment

lbs ft ft-lbs lbs ft ft-lbs

STRUCTURES 38381 2683580 EQUIPMENT 23703 1551014

 Wing 14201 74.8 1062591 Flight Controls 2819 70.0 197323

 Retracting Canard 1130 30.0 33888 Instruments 159 10.0 1593

0 Hydraulics 1109 80.0 88720

 Chin Rudder 207 6.7 1387 Electrical 2126 80.0 170088

 Fuselage 14818 60.0 889080 Avionics 1387 10.0 13870

0 Air Conditioning 1512 60.0 90714

 Nacelle & Inlet 1100 112.9 124190 Anti-Icing 1307 75.0 97988

 Nacelle pylon 2656 110.0 292160 APU installed 660 115.0 75900

0 Furnishings & Handling 7843 60.0 470586

 Main Landing Gear 3488 77.0 268568 Other & misc 4781 72.0 344232

 Nose Landing Gear 781 15.0 11717 5% Empty Weight Allowance 3751 75.9 284744

PROPULSION 12929 1460290 TOTAL WEIGHT EMPTY 78764 75.9 5979628

 Engine 5674 112.9 640595

 Open Rotor & Mech 5038 118.0 594484 USEFUL LOAD 95436

 Accessories/Mounts 1708 110.0 187880 Crew 400 9.0 3600

 Engine Controls 42 110.0 4620 Fuel 51856 75.0 3889219

0 Oil 110 112.0 12342

0 Passengers 29160 65.0 1895400

0 Payload 13910 65.0 904150

0

 Fuel System 467 70.0 32711 TAKEOFF GROSS WEIGHT 174200 72.8 12684339
NASA/TM—2011-217130 33 

TABLE 5.—NAS2 SIZED WEIGHT STATEMENT 

Weight Loc Moment Weight Loc Moment

lbs ft ft-lbs lbs ft ft-lbs

STRUCTURES 28094 1956000 EQUIPMENT 23024 1501234

 Wing 8974 74.8 671463 Flight Controls 2634 70.0 184373

 Retracting Canard 659 30.0 19773 Instruments 152 10.0 1524

0 Hydraulics 1021 80.0 81688

 Chin Rudder 170 6.7 1141 Electrical 2126 80.0 170088

 Fuselage 12431 60.0 745842 Avionics 1387 10.0 13870

0 Air Conditioning 1512 60.0 90714

 Nacelle & Inlet 747 112.9 84380 Anti-Icing 919 75.0 68955

 Nacelle pylon 2276 110.0 250381 APU installed 660 115.0 75900

0 Furnishings & Handling 7832 60.0 469890

 Main Landing Gear 2266 77.0 174459 Other & misc 4781 72.0 344232

 Nose Landing Gear 571 15.0 8561 5% Empty Weight Allowance 3000 74.3 222815

PROPULSION 8875 999066 TOTAL WEIGHT EMPTY 62993 74.3 4679116

 Engine 3855 112.9 435250

 Open Rotor & Mech 3423 118.0 403920 USEFUL LOAD 59599

 Accessories/Mounts 1160 110.0 127654 Crew 400 9.0 3600

 Engine Controls 42 110.0 4620 Fuel 16029 75.0 1202153

0 Oil 100 112.0 11200

0 Passengers 29160 65.0 1895400

0 Payload 13910 65.0 904150

0

 Fuel System 395 70.0 27622 TAKEOFF GROSS WEIGHT 122592 70.9 8695619

With all these assumptions, reductions, penalties, and margins, the unsized NAS2 design empty 

weight estimate of 78,764 is 14 percent less than the B-737. While in the right direction, this is far less 

than the top-level goal of a 30 percent reduction. Since Boeing itself claims a 20 to 25 percent weight 

savings on the current-technology B-787 versus the same design using conventional construction as on 

the B-737, it is likely that these weight results are conservative for the N+3 timeframe. However, in the 

absence of detailed structural calculations and subsystems engineering, such conservatism seems 

appropriate. 

Note that the Center of Gravity (cg) travel is a bit excessive, and needs further attention in a 

subsequent design effort. This is within the “normal” band of early layout problems, and is nothing to be 

worried about. 

The following weight statement is for the resized NAS2 as described in Table 5. Component weights 

are reduced because the wings, canard, tires, engines, and other items are smaller as a result of the 

“leverage” effect of the sizing process. 

11.4 Propulsion 

The Open Rotor propulsion system was defined by NASA-GRC personnel for this study. Its analysis 

assumed an N+3 collection of propulsion technologies advances in addition to the Open Rotor itself, 

including an advanced core with an Overall Pressure Ratio (OPR) of 70, advanced materials, improved 

cooling schemes, and greater component efficiencies. Note that these technologies, when applied to a 

more-traditional advanced high-bypass turbofan, would yield much of the expected benefit and probably 

at a reduced weight. Considering the potential noise problems of the Open Rotor, this would make an 

excellent trade study. 

 
NASA/TM—2011-217130 34 

Data representing NASA’s Open Rotor engine was provided as a collection of thrust and fuel 

consumption at various speeds, altitudes, and atmospheric conditions. To populate a complete engine 

map, those values were extrapolated to other conditions based on previous engine data and in some cases, 

“eyeball” fairing. The resulting data curves, shown in Figures 20 and 21, were reviewed by NASA-GRC 

personnel and found to be reasonable for a first-pass design study. The circled points are those originally 

supplied by NASA. 

The engine was assumed to be “rubber”, i.e., scalable to any desired thrust level. For this study, engine 

scaling was used to provide the same thrust-to-weight ratio (T/W) as the B-737. Standard engine scaling 

rules were applied when the NAS2 design was sized to the mission, reducing the thrust requirement. 

Figure 20.—Open rotor engine thrust. 

Figure 21.—Open rotor engine fuel consumption. 
NASA/TM—2011-217130 35 

11.5 Sizing 

The NAS2 design was then resized, based on all the calculations above, to the same mission as the 

baseline B-737. Since the drag, fuel consumption, and empty weight are all much better than the B-737, a 

smaller and lighter airplane can carry the same passengers and payload over the same distance. Of course, 

the fuselage has to be about the same size to hold the people, but everything else can shrink. 

Sizing results are shown in Table 6. The summed fuel consumption is in the far-right column, with 

the total fuel weight, empty weight, and aircraft takeoff gross weight at the bottom. For a full explanation 

of these terms and methods, see Raymer (Ref. 2). 

TABLE 6.—NAS2 SIZING CALCULATIONS 

 MISSION SEGMENT Wi/WO FUEL BURN (lbs-m) 

 WEIGHT FRACTION SEGMENT TOTAL 

 1 TAKEOFF SEGMENT 0.9989 0.9989 143.8 143.8 

 2 TAKEOFF SEGMENT 0.9987 0.9976 166.7 310.4 

 3 CLIMB and/or ACCEL. 0.9857 0.9833 1809.4 2119.8 

 4 CRUISE SEGMENT 0.9472 0.9314 6607.4 8727.3 

 5 CLIMB and/or ACCEL. 0.9994 0.9308 68.2 8795.5 

 6 CRUISE SEGMENT 0.9481 0.8825 6149.3 14944.8 

 7 DESCENT ANALYSIS 0.9976 0.8804 269.3 15214.1 

 8 CLIMB and/or ACCEL. 0.9905 0.8720 1058.6 16272.7 

 9 CRUISE SEGMENT 0.9951 0.8678 542.2 16815.0 

 10 DESCENT SEGMENT 0.9950 0.8634 551.8 17366.7 

 11 LANDING SEGMENT 0.9950 0.8591 549.0 17915.7 

 Reserve & trap : 1074.9 

 Total fuel : 18990.7 

 Seg. 4 CRUISE : 410.6 kts at 34000.0 ft RANGE = 1387.0 nm 

 Seg. 6 CRUISE : 410.6 kts at 34000.0 ft RANGE = 1387.0 nm 

 Seg. 9 CRUISE : 370.6 kts at 25000.0 ft RANGE = 200.0 nm 

 FUEL WEIGHT = 18991.2 EMPTY WEIGHT = 64607.6 

 USEFUL LOAD (less Wf)= 43570.0 AIRCRAFT GROSS WEIGHT = 127168.8 

The results are encouraging. The sized takeoff gross weight of the NAS2, carrying the same 

passengers and payload over the same mission range, is 127,169 lb, burning 18,991 lb of fuel. This is 

60 percent less fuel consumption that the B-737-800 which weighs 174,200 lb and burns 46,750 lb of 

fuel. While our goal was a 70 percent reduction, this 60 percent reduction is quite an accomplishment and, 

if the proposed design and technologies all work as hoped, has a realistic chance of success. 

The NAS2 design was redrawn to reflect the sizing results, with wings, engines, canard, and landing 

gear reduced in size. The design layouts shown above all reflect this resized version. 
NASA/TM—2011-217130 36 

11.6 Issues and Way-Forward 

The NAS2 design concept is exciting and seems to offer a real path forward to obtain a tremendous 

saving in fuel consumption. However, this study was done with a limited budget so certain issues were 

not addressed. Some of these are 

 Noise reduction is a key issue for N+3, and the selected Open Rotor propulsion concept is 

inherently noisier than a traditional turbofan. Several of the initial design concepts in this study 

used either the wing or the horizontal tail to prevent engine noise from reaching the ground. In the 

interest of aerodynamic efficiency, the final preferred concept Tailless Open Rotor does not have 

such noise blocking geometry, so the open rotor itself must be noise-suppressed. Failing that, it 

may be necessary to put a shroud around the rotor, turning the engine into an ultra-high-bypass 

turbofan. 

 The Open Rotor provides no natural ballistic containment in the event of rotor or other engine 

failure. The nacelles are positioned behind the fuselage pressure vessel so its penetration is not a 

problem, but one engine failure could damage the other engine. It may prove necessary to offset 

the engines or provide a stub vertical tail as a “wall” between them, or reposition them 

downwards so that the aft fuselage separates them. Again, a shroud could be used around the 

rotor, making the engine a turbofan. 

 The thrust axis seems rather high and may pose problems in nose-wheel liftoff and perhaps in 

cruise trim. The canard elevator/flap was deliberately oversized to handle the nose-wheel liftoff 

problem, but detailed analysis is required. If necessary there is some room to move the nacelles 

downwards, but this would bring the rotors close to the ground in an extreme nose-up takeoff or 

landing. 

 If the canard should fail to deploy for landing, the no-canard, no-flaps landing speed may be 

excessive even for emergency operation at extra-long runways. If true, either the canard extension 

needs some failsafe emergency actuation, or the design may need modification to provide a safe 

landing in its absence. 

 An aircraft which is unstable in yaw may, in a worst-case scenario, experience an extreme yaw 

upset in excess of available recovery control power. This must be determined to occur at such a 

low probability that the failure mode likelihood is in line with other catastrophes such as wing 

structural failure, total loss of control actuation, etc. 

 The flexibility of the Active Aeroelastic Wing will result in greater bending and torsional 

deflections during its life, leading to greater fatigue potential. This may limit the permissible 

weight savings, and will require sophisticated health monitoring systems to ensure that 

accumulated damage is detectible. 

 
NASA/TM—2011-217130 37 

This study was performed on a limited budget by a small company. The answers are good enough and 

interesting enough that further study appears warranted, but nobody should claim that these answers are 

final or that anything has been “proven.” Areas recommended for further study include: 

 Acquisition and use of a detailed Open Rotor engine model including full thrust and fuel flow 

calculations, geometry, installation guidance, etc. 

 Trade studies of Open Rotor versus Ultra High Bypass Turbofan 

 Computational aerodynamics (CFD) for analysis and design shape optimization 

 Structural finite element analysis (FEM) and advanced materials trade studies 

 Aeroelastic 6-DOF dynamic simulation including the use of the Active Aeroelastic Wing 

technology and the controllability issues in pitch and yaw 

 Weights estimates based on structural and subsystem design and analysis 

 Numerous trade studies of propulsion, materials, subsystems, and design configuration 

 Multidisciplinary Design Optimization (MDO) studies 

 Detailed performance analysis including FAA climb specifications, takeoff, etc. 

 Additional design iterations with detailed CAD work to verify installation of the retracting 

canard, landing gear kinematics and actuation, fuel tanks, passenger compartment, cargo bays and 

handling, and nacelle/pylon design. 

 Development planning and cost analysis 

 
NASA/TM—2011-217130 38 

12.0 Summary and Conclusions 

Conceptual Research Corporation has completed its N+3-related study of advanced transport aircraft 

in support of NASA’s ongoing Subsonic Fixed Wing project. The overall goal was to assess what it 

would take to obtain a 70 percent reduction in fuel use for a 2035 timeframe commercial airliner, and to 

define vehicle concepts and technologies that show promise in that direction. 

Parametric studies using the Boeing 737 found that a simultaneous 30 percent improvement in vehicle 

weight, aerodynamic drag (parasitic and induced), and engine fuel consumption will together add up to 

the desired 70 percent fuel savings. To obtain such improvements, a design approach with reduced wetted 

surface area and increased laminar flow is essential. To reduce drag-due-to-lift, simple physics says that 

the aspect ratio must be substantially increased. 

The Active Aeroelastic Wing technology, flight tested and proven, is suggested to permit this with 

minimal weight penalty. It also facilitates the tailless design approach because it allows the use of the 

wing for pitch, yaw, and maneuver load control as well as the obvious roll control. 

Figure 22.—NAS2 on the ramp. 
NASA/TM—2011-217130 39 

A propulsion approach which increases the effective bypass ratio is needed, with the Open Rotor as 

the most-likely emerging concept to obtain this. 

Aircraft weight reductions are required across-the-board since no one weight category has enough 

weight to reach the goal alone. 

Notional design studies of various novel configuration arrangements showed several concepts with 

potential for benefit. A top-level evaluation based on trimmed maximum lift coefficient and liftnormalized wetted surface area was used to rank these concepts. From this it was concluded that the 

designs which eliminated the tails showed the greatest promise. Note that this result is specific to smaller 

airliners. For the largest airliners, it is likely that the Blended Wing Body would show higher promise, but 

for a smaller airliner the BWB suffers from internal packaging inefficiencies. 

The CRC-developed NAS2 Tailless Open Rotor design concept, sized to the same payload and range 

as the B-737-800, was found to burn 60 percent less fuel over the same mission. While a lot of work is 

required to mature and validate this design concept, it offers a way-forward towards the N+3 goals and, 

while admittedly exotic, is based on proven or clearly-emerging technologies and design approaches. It 

should work. 

Figure 23.—NAS2 three-view. 

NASA/TM—2011-217130 41 

Appendix—Parametric Sizing Data 

TABLE 7.—PARAMETRIC SIZING FOR 4 PARAMETER BOX-BEHNKEN SCHEME RESPONSE SURFACE 

 Cd0 K sfc W

Run Wo (lbs) We (lbs) Wf (lbs)

0 D0 K0 C0 W0 F0 174361 92048 46455

1 D0 K0 C0 -25% W0-25% F1 108417 49820 23140

2 D0-50% K0-50% C0 -25% W0-25% F2 92594 44706 12430

3 D0 K0-50% C0 -25% W0-25% F3 100313 47232 17622

4 D0-50% K0 C0 -25% W0-25% F4 97993 46480 16055

5 D0 K0-25% C0 W0-25% F5 117809 52743 29607

6 D0-50% K0-25% C0 W0-25% F6 103768 48343 19967

7 D0 K0-25% C0-50% W0-25% F7 93393 44971 12964

8 D0-50% K0-25% C0-50% W0-25% F8 87853 43122 9273

9 D0 K0-25% C0 -25% W0 F9 144215 80799 27958

10 D0-50% K0-25% C0 -25% W0 F10 130162 75307 19397

11 D0 K0-25% C0 -25% W0-50% F11 77192 26304 15430

12 D0-50% K0-25% C0 -25% W0-50% F12 71127 24867 10802

13 D0-25% K0 C0 W0-25% F13 115468 52021 27988

14 D0-25% K0-50% C0 W0-25% F14 105030 48746 20826

15 D0-25% K0 C0 - 50% W0-25% F15 92413 44646 12308

16 D0-25% K0-50% C0 - 50% W0-25% F16 88337 43285 9594

17 D0-25% K0 C0-25% W0 F17 141860 79891 26511

18 D0-25% K0-50% C0-25% W0 F18 131182 75712 20012

19 D0-25% K0 C0-25% W0-50% F19 76086 26045 14583

20 D0-25% K0-50% C0-25% W0-50% F20 71697 25003 11235

21 D0-25% K0 - 25% C0 - 50% W0 F21 122266 72140 14667

22 D0-25% K0 - 25% C0 W0 F22 153280 84253 33569

23 D0-25% K0 - 25% C0 - 50% W0-50% F23 67937 24096 8383

24 D0-25% K0 - 25% C0 W0-50% F24 80675 27114 18104

25 D0-25% K0 - 25% C0 -25% W0 - 25% F25 99819 47072 17288

Base D0 K0 C0 W0 F0 174361 92048 46455

D0-50% K0 C0 W0 F26 149297 82743 31096

 D0 K0-50% C0 W0 F27 154288 84632 34196

 D0 K0 C0-50% W0 F28 130187 75317 19412

 D0 K0 C0 W0-50% F29 88958 28995 24505

 Cd0 K sfc W Wo (lbs) We (lbs) Wf (lbs)

1-Var

Results - true scaling
NASA/TM—2011-217130 42 

References 

1. anon., Advanced Concepts Studies Awardees, 

www.aeronautics.nasa.gov/nra_awardees_10_06_08.htm, 2008 

2. Raymer, D., AIRCRAFT DESIGN: A Conceptual Approach, American Institute of Aeronautics and 

Astronautics, Washington, D.C., Fourth Edition 2006. 

3. Hoak, D., Ellison, D., et al., USAF DATCOM,Air Force Flight Dynamics Lab., Wright-Patterson 

AFB, OH. 

4. Staton, R., Cargo/Transport Statistical Weight Estimation Equations,Vought Aircraft Rept. 2-59320 

9R-50549, 1969. 

5. Raymer, D., “NASA Advanced Technology Subsonic Transport Study Final Report Briefing” 

(NASA-GRC Contract), Conceptual Research Corporation, Los Angeles, CA, Oct. 2009. 

6. Raymer, D., “NASA Advanced Technology Subsonic Transport Study Phase 2 Final Report 

Briefing” (NASA-GRC Contract), Conceptual Research Corporation, Los Angeles, CA, Oct. 2010. 

7. Boeing Commercial Airplanes, “737 Airplane Characteristics For Airport Planning,” Publication D6-

58325-6, October 2005. 

8. CFM56 Engine Website, URL: http://www.cfm56.com, January 2009. 

9. Perkins, D., Wilson, J., and Raymer, D., “An Evaluation of Performance Metrics for High Efficiency 

Tube-and-Wing Aircraft Entering Service in 2030-35,” NASA/TM—2011-217264. 

10. Wolkovitch, J., “The Joined Wing: An Overview,” AIAA Journal of Aircraft, March 1986. 

11. Croft, J., “NASA, Boeing complete X-48B flight test campaign,” Flight International, 06/04/10. 

12. Raymer, D., Living In The Future: The Education and Adventures of an Advanced Aircraft Designer, 

Design Dimension Press, Los Angeles, CA, 2009. 

13. Huber, P., and Schellenger, H., “X-31 Quasi-Tailless Flight Demonstration,” NASA N95-14243, 

1995. 

14. Phillips, D., “X-36 Tailless Prototype Fighter Completes Phase 1 Flight Testing,” Press Release, 

McDonnell Douglas Corp., 1997. 

15. Active Flexible Wing Aircraft Control System, US Patent 5082207, 1/21/1992. 

16. anon., “Boeing X-53 Active Aeroelastic Wing,” Wikipedia, 12-2010. 

17. Raymer, D., Aircraft Aerodynamic Analysis on a Personal Computer, SAE Paper 932530, Aerotech 

93, Sept 27, 1993. 

18. Rizzi, A., and Zhang, M., “Modeling and Simulating Aircraft Stability & Control (the SimSAC 

Project),” AIAA Guidance, Navigation, and Control Conference, Toronto, Canada, 2010. 

19. Rizzi, A., Oppelstrup, J., Zhang, M. and Tomac, M., “Coupling Parametric Aircraft Lofting to CFD & 

CSM Grid Generation for Conceptual Design,” AIAA Aerospace Sciences Meeting, Orlando, FL, 

2011. 
REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-11-2011 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Advanced Technology Subsonic Transport Study 

N+3 Technologies and Design Concepts

5a. CONTRACT NUMBER

NNC08VF71P 

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Raymer, Daniel, P.; Wilson, Jack; Perkins, H., Douglas; Rizzi, Arthur; Zhang, Mengmeng; 

Puentes, Alfredo, Ramirez 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 561581.02.08.03.13.05 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-17819 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2011-217130 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Category: 05 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Conceptual Research Corporation, the Science of the Possible, has completed a two-year study of concepts and technologies for future airliners in the 180-passenger class. This 

NASA-funded contract was primarily focused on the ambitious goal of a 70 percent reduction in fuel consumption versus the market-dominating Boeing 737-800. The study is 

related to the “N+3” contracts awarded in 2008 by NASA’s Aeronautics Research Mission Directorate to teams led by Boeing, GE Aviation, MIT, and Northrop Grumman, but 

with more modest goals and funding. CRC’s contract featured a predominant emphasis on propulsion and fuel consumption, but since fuel consumption depends upon air 

vehicle design as much as on propulsion technology, the study included notional vehicle design, analysis, and parametric studies. Other NASA goals including NOx and noise 

reduction are of long-standing interest but were not highlighted in this study, other than their inclusion in the propulsion system provided to CRC by NASA. The B-737-800 

was used as a benchmark, parametric tool, and design point of departure. It was modeled in the RDS-Professional aircraft design software then subjected to extensive 

parametric variations of parasitic drag, drag-due-to-lift, specific fuel consumption, and unsized empty weight. These studies indicated that the goal of a 70 percent reduction in 

fuel consumption could be attained with roughly a 30 percent improvement in all four parameters. The results were then fit to a Response Surface and coded for ease of use in 

subsequent trade studies. Potential technologies to obtain such savings were identified and discussed. More than 16 advanced concept designs were then prepared, attempting 

to investigate almost every possible emerging concept for application to this class airliner. A preliminary assessment of these concepts was done based on their total wetted 

area after design normalization of trimmed maximum lift. This assessment points towards a Tailless Airliner concept which was designed and analyzed in some detail. To make 

it work, a retracting canard and an all-moving chin rudder were employed, along with the use of the Active Aeroelastic Wing technology. Results indicate that a 60 percent 

savings in fuel burn may be credibly attained, but this depends upon a lot of technology maturation, concept development, and risk reduction. This should be expected-such a 

dramatic reduction in fuel consumption is a “game changer” in the world of commercial aviation. It won’t be easy.

15. SUBJECT TERMS

Passenger aircraft; Subsonic aircraft 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

49 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18