Sam Bruner, Scott Baber, Chris Harris, Nicholas Caldwell, Peter Keding, Kyle Rahrig, and Luck Pho

Northrop Grumman Corporation, El Segundo, California

Richard Wlezian

Tufts University, Medford, Massachusetts

NASA N+3 Subsonic Fixed Wing Silent Effi cient 

Low-Emissions Commercial Transport (SELECT) 

Vehicle Study

Revision A

NASA/CR—2010-216798

November 2010
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 NASA Center for AeroSpace Information (CASI)

 7115 Standard Drive

 Hanover, MD 21076–1320
Sam Bruner, Scott Baber, Chris Harris, Nicholas Caldwell, Peter Keding, Kyle Rahrig, and Luck Pho

Northrop Grumman Corporation, El Segundo, California

Richard Wlezian

Tufts University, Medford, Massachusetts

NASA N+3 Subsonic Fixed Wing Silent Effi cient 

Low-Emissions Commercial Transport (SELECT) 

Vehicle Study

Revision A

NASA/CR—2010-216798

November 2010

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC08CA86C
Acknowledgments

The authors would like to thank Doug Perkins, our NASA Contracting Offi ce Technical Representative (COTR), and Brian Allan, NASA Langley, 

for their support and guidance during the execution of our study. Their feedback was particularly helpful in keeping us on target during the 

preparation of this fi nal report. We would like to thank Arnie McCullers, NASA Langley, for his dedication in timely response to our questions 

on FLOPS, even when dealing with his own medical issues. This work was funded by the Subsonic Fixed Wing project, of the Fundamental 

Aeronautics Program, under NASA contract NNC08CA86C. Contributors: The following individuals composed the core technical team that 

performed this study over the 18-month period from October 2008 through March of 2010. The contractor and subcontractors were generous in 

providing numerous other individuals who were integral to the support of very specialized aspects of this effort. Northrop Grumman: Scott Collins 

provided programmatic support and guidance and technical review. Dr. Sam Bruner was the Principal Investigator and Program Manager. Chris 

Harris was the Systems Analysis IPT Lead, and supported technology selection and planning, propulsion integration, the scenario evaluation, 

and the requirements study. Nicholas Caldwell supported propulsion integration, technology modeling, and vehicle trades/optimization, and 

led the editing of the fi nal report. Peter Keding supported technology modeling, aerodynamics and performance evaluation, and vehicle trades/

optimization. Scott Baber supported confi guration design and vehicle trades/optimization. Luck Pho supported aerodynamic design and 

performance. Kyle Rahrig performed acoustic modeling and analysis. Rolls-Royce Liberty Works: David Eames led the RRLW propulsion 

support and advanced engine architecture design, and propulsion technologies study. Tufts University: Dr. Rich Wlezien performed the future 

scenario research and analysis, and led the technology maturation planning effort. Spirit Aerosystems: Dr. Judy Gallman led the technology 

evaluation of acoustic airframe treatment technologies, and the design of an advanced acoustic inlet liner. Sensis: Dave Miller led the effort to 

determine N+3 requirements, as well as perform N+3 simulations in the NAS.

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://gltrs.grc.nasa.gov

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 NASA technical management.
	



 


 


 

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1 EXECUTIVE SUMMARY 

Northrop Grumman has developed an advanced aircraft concept whose mission capabilities enable it 

to fill a broad, primary need within the future scenario envisioned for the 2030-2035 timeframe. This 

effort was conducted under NASA Contract NNC08CA86C. No single vehicle type is adequate to meet 

all needs simultaneously. Multi-objective studies identified vehicle concepts with the best potential for 

meeting the combined goals. Trade studies provided insight into individual goals, in particular as they 

related to the future scenario for the 2030-35 and beyond timeframe. It is possible that future capacity 

challenges at major airports can be addressed by increased use of exiting runways and reliever airports, 

without the need for public investment in capital infrastructure and without the need for short take-off and 

landing (STOL) capable aircraft. 

The primary vehicle requirements that drove technology selection, configuration selection, and 

optimization were derived from the published NASA N+3 goals shown in Table 1-1. These goals were 

further refined based on an overarching future scenario study that captured likely future economic and 

operational factors. 

Table 1-1. NASA Generational Aircraft Goals 

A study of the current vehicle fleet mix was used to determine the most broadly applicable airframe in 

use today based on stage-lengths served, and then extrapolate that based on the future scenarios. 

Simulations of the future vehicle in the National Airspace System (NAS) with the implementation of 

NextGen Air Traffic Management (ATM) capabilities were used to assess the impact of vehicle sizing 

and metroplex operations on projected traffic levels. The ability to adopt metroplex-type operations was 

assessed by a study of runways at existing airports, combined with NAS simulations. It was found that the 

use of reliever airports with at least one runway of 5,000 ft length within 70 nautical miles of major 

population centers would enable metroplex operations while keeping pace with projected traffic levels. A 

single set of design requirements emerged from these studies, as summarized in Table 1-2. 

Table 1-2. Mission Requirements 

Several other critical design concepts emerged. First, it became evident that cruise Mach number can 

be reduced to save fuel burn without significant impact on throughput or delay in the overall air 

transportation system. A byproduct of cruising slower is that the airplane will fly higher to achieve the 

desired cruise lift coefficient for best efficiency. Second, short takeoff and landing (STOL) and extremely 

Corners of the 

Trade Space

N+1 (2005 Entry Into Service)

Generation

Conventional Tube and Wing

(Relative to B737/CFM56)

N+2 (2020 Initial Operating 

Capability) Generation

Unconventional

Hybrid Wing Body

(Relative to B777/GE90)

-32 dB -42 dB -71 dB

N+3 (2030

Advanced Aircraft Concepts

(Relative to User -Defined 

Reference)

-60% -75% Better than -75%

Performance: Aircraft 

Fuel Burn

-33% -40% Better Than -70%

Noise

(Cum Below Stage 4)

LTO NOx Emissions

(Below CAEP 6)

Performance: Field 

Length -33% -50% Exploit Metroplex Concepts

- 2035 Service Entry)

Mission Requirements

Range (with reserves): 1600 nm

Passengers: 120

Balanced Field Length (Sea Level/Standard Day): 5000 ft

Landing Distance (Sea Level/Standard Day): 5000 ft

Minimum Cruise Mach: 0.75

-71 dB

Better than -75%

Better Than -70%

Exploit Metroplex Concepts

NASA/CR—2010-216798 1
short takeoff and landing (ESTOL) capability is not required to meet the 5000 foot runway requirement. 

Relaxing the design space on cruise Mach number and maximum lift coefficient provided more latitude 

for achieving desired improvements in fuel burn. Finally, a wing span limitation of 150 feet was imposed 

to accommodate space restrictions at metroplex airports. 

Two important tasks were undertaken before beginning aircraft design studies. The Flight 

Optimization System (FLOPS) vehicle sizing program developed by NASA was to be the engine for all 

optimization studies. First, FLOPS was calibrated using a 737-800 with CFM56-7B27 engines to match 

published weight and performance data to within an accepted tolerance. Next, a reference vehicle was 

derived as a baseline against which performance and noise metrics could be measured. This reference 

vehicle was derived from a 737-500, resized to match the requirements of Table 1-2 (new wing size, new 

thrust level, etc.), but with the same level of technology as the 737. This reference vehicle would be the 

“yardstick” for measuring improvement in subsequent designs. 

A wide design space of airframe configurations was initially considered. Technologies were then 

systematically applied in a conceptual design process. Low-order models and a mix of empirical and 

analytic estimation methods were used to describe component-level benefits over the vehicle flight 

envelope. Weighting factors for the N+3 goals were developed as part of the future scenario analysis, and 

these weighting factors were incorporated into a System Effectiveness Rating (SER). The aircraft with the 

highest SER emerged as the preferred concept vehicle. The preferred concept, summarized in Figure 1-1, 

reflects a NextGen-capable platform that best meets the potential future traffic demands and satisfies the 

N+3 goals and the mission requirements listed in Table 1-3. 

Figure 1-1. N+3 Preferred Configuration 3-View and Mission Performance 

Range [nmi] PAX Max TOGW [lb] Cruise Alt. [ft] Cruise Mach Cruise SFC [pph/lbf] BFL [ft]

1600 120 80478 45,000 0.75 0.451 4999

N+3 Mission & Performance

Technology Suite

Three-Shaft Turbofan Engine

- Ultra-High Bypass Ratio of ~18

- Compressor Intercooling

- Lean-Burn CMC Combustor

- CMC Turbine Blades

- Cooled Cooling Air Turbine

- Memory Metal Nozzle Design

- Porous Ceramic Nozzle Material

- Endothermic Fuel System

- Variable Geometry Nozzle

Swept Wing Laminar Flow

Large Integrated Structures

Aeroservoelastic Structures

M5 Ultra High-Performance Fiber

Carbon Nanotube Electrical Cables

3-D Woven and Stitched Composites

Advanced Metallics

Landing Gear Fairings

Advanced Acoustic Inlet Liner

Technology Suite

Three-Shaft Turbofan Engine

-Ultra-High Bypass Ratio of ~18

-CMC Turbine Blades

-Lean-Burn CMC Combustor

-Intercooled Compressor Stages

-Swept Fan Outlet Guide Vanes

-Fan Blade Sweep Design

-Lifhgtweight Fan/Fan Cowl

-Compressor Flow Control

-Active Compressor Clearance Control

-Variable Geometry Nozzles

Swept Wing Laminar Flow

Large Integrated Structures

Aeroservoelastic Structures

M5 Ultra High-Performance Fiber

Carbon Nanotube Electrical Cables

3-D Woven and Stitched Composites

Advanced Metallics

Landing Gear Fairings

Advanced Acoustic Inlet Liner

NASA/CR—2010-216798 2
Table 1-3. N+3 Goals Achievement 

Several key technologies influence the outcome of the design. Advanced technology engines 

contribute to improved fuel efficiency, noise, and nitrous oxide (NOx) emissions. Advanced materials 

reduce the empty weight of the aircraft and enable high aspect ratio, highly efficient, wings. Swept-wing 

laminar flow reduces drag and further improves cruise efficiency. Several technologies that initially 

looked promising did not prove their worth after further study. Hybrid wing-body configurations did not 

have the volumetric packing factor to be viable in this class of vehicle. Open rotor engine architectures 

provide significant opportunity to improve SFC, but the extra weight associated with the gear box and 

propulsor rotors offset much of that benefit in this class of vehicle. The lack of maturity in noise 

prediction methods for open-rotor configurations, and the associated uncertainty with community noise 

estimates, also contributed to this technology being set aside. However, the studies did imply that hybrid 

wing-bodies and open rotors might show benefits in application to extremely large long-range aircraft. A 

different set of mission requirements might have resulted in a different preferred technology set applied to 

a different configuration. 

A key product of this study was the development of technology maturation plans (TMPs). These 

TMPs will assist NASA in planning its future investment in a manner that supports enabling critical 

technology items at Technology Readiness Level (TRL) 6 by the 2025 timeframe. This is essential to 

support an entry into service (EIS) date of 2030-2035. Technologies and architectures that strongly 

influence attaining critical mission performance metrics must be at a solid TRL 6 or better in order to 

minimize commercial risk. TMPs for the selected technologies are available in Section 7. Technologies 

that were considered but not included for further consideration are listed in Appendix B, including the 

reasoning for their dismissal from the preferred technology set. In many cases, sufficient TRL in the 

required timeframe was a determining factor. In others cases, incompatibility with competing 

technologies played the greatest role in the exclusion of certain technologies from the final package. 

This study shows that significant advances in efficiency, community noise, and NOx emissions are 

attainable with focused technology advances in propulsion systems, aerodynamic technology, material 

systems, and other technologies. Furthermore, the system-level achievements are shown to be consistent 

with the broad needs identified in the scenario studies for the 2030-2035 timeframe. Wise investment by 

NASA in these enabling technologies will benefit not only the flying public, but the environment as well. 

Metric Overall Scenario Weighting Factor

Noise -71 EPNdB relative to Stage 4 200.3 EPNdB 201.7 EPNdB 69.6 EPNdB below Stage 4 20.00%

Fuel Burn 70% below reference vehicle 7514 lbf 9145 lbf 63.49% below reference vehicle 33.33%

Emissions 75% below CAEP/6 25.7 g/kN 9.69 g/kN 90.59% below CAEP/6 33.33%

Balanced Field Length exploit metroplex 5,000 ft 4,999 ft 13.33%

Requirement Actual Achievement

exploited metroplex

NASA/CR—2010-216798 3
2 TECHNICAL APPROACH 

Successfully identifying our preferred air vehicle concept for the program's objectives began with 

defining the anticipated future operating scenario. Through evaluating operations in the context of 

vehicle performance, passenger count, turnaround time, requirements were determined and then flowed to 

the design team. This was accomplished by investigating currently available data used in prior NASA 

studies, by performing new simulations of the NAS using an approximate model of an N+3 vehicle, and 

by assessing the available infrastructure and how it is impacted by NextGen ATM implementation, which 

is assumed to be fully completed by the EIS date. Figure 2-1 shows a summarized description of the 

technical approach used in this program. 

Figure 2-1. Technical Approach 

Requirements and weighting factors for the N+3 goals were also derived from examining a set of 

scenarios for the N+3 timeframe. The scenarios project the future world political, social, and economic 

states with a focus on the United States. A set of fifteen published reports was used to compile and 

analyze the drivers most relevant to aviation, including energy resource supply and demand, political 

turmoil or order, and social protest to environmental factors. These drivers were grouped by common 

macroeconomic effects and were then used to develop an overall scenario that could bound estimates on 

future traffic levels, and quantify the relative importance of the N+3 goals. 

Modeling tools to predict performance and weight were calibrated using a Boeing 737-800. 

Comparisons of other NGC tools were used to further validate and assess the accuracy of system weights, 

acoustic predictions, and aerodynamic performance. 

Subsection 6.1

Trade Studies

Subsection 5.3

Technology Assessment

Subsection 5.2

Preliminary Concepts

Section 6

Analysis

Subsection 6.4

Exit Criteria

Section 7 Technology Risk

Phase 2

Technology 

Roadmaps, Findings 

and Recommendations 

Phase 2 Proposal and 

Final Report

Pha

Subsection 3.11 &6.5

NASA/CR—2010-216798 4
In order to specifically measure the N+3 metrics with respect to a state-of-the-art commercial 

transport, a reference vehicle was required. The Boeing 737-500 with CFM56-3B1 engines was resized to 

meet the N+3 mission requirements, while holding a constant level of technology. The resulting reference 

vehicle had a larger wing area and slightly more thrust than the 737-500. This reference vehicle served as 

a baseline from which the N+3 fuel burn goal was defined. Similarly, this vehicle was used as a valuable 

reference for weight, emissions, and acoustic results. 

The next task was to develop a super-set of configurations and technologies that best applied to the 

vehicle mission and requirements defined by the NAS analyses and scenario study. For the technology 

application, a quality function deployment (QFD) method was used to identify a preliminary set of 

technologies best suited to the future mission requirements. A technology database, with input from all 

disciplines, allowed the ranking of a large number of candidate technologies to quickly and effectively 

identify the best suites for a particular configuration. Each technology was also assessed for compatibility 

with other technologies, to improve the quality of the identified sets in terms of vehicle-level impact. Risk 

level was also factored into the ranking process to apply an up-front, first-order risk reduction to the 

ranking. Technologies that were deemed too immature for the N+3 timeframe were discarded. Figure 2-2 

shows the technological development and implementation process that was assumed as a driver behind 

technology selection. Total development time from concept to EIS must allow key technologies that 

fundamentally influence aircraft size and configuration to be significantly mature to manage commercial 

risk. With EIS in 2030-2035, a technology would require substantial maturity, TRL 6, by 2025. This point 

defines the minimum level of acceptable commercial risk prior to initiating the design and development 

phase of the aircraft. After this point, technology maturation continues, and risk is further reduced. 

Subsequent trade studies and analysis would later narrow this technology list further before arriving at a 

preferred technology set. 

Figure 2-2. Assumed Technology Maturation and Implementation Timeframe 

NASA/CR—2010-216798 5
During the conceptual design and examination of the airframe configurations, aerodynamic models 

were implemented using a mix of empirical and analytical data. Technology packages were assembled to 

maximize the effectiveness of each airframe against the N+3 goals. A reduced set of candidate 

technologies was eventually identified that best met all of the N+3 goals on the basis of the technology 

effectiveness rating (TER), an all-encompassing parameter that is a function of the N+3 goals. These 

technologies were then modeled using various low-order techniques to simulate their performance and 

weight impacts across the vehicle flight envelope. Trades were performed with and without the 

technology packages, since many of the technology applications (for example, laminar flow) required 

trades to other configuration design parameters. Models for the configurations deemed applicable were 

refined further in a CAD environment that allowed more technical fidelity. Advanced engine models were 

developed for the appropriate thrust class for these vehicles, resulting in models for both a high bypass 

ratio three-shaft turbofan and an open rotor engine architectures. Both engine concepts were investigated 

in various installations, including aft-mounted, embedded, and under-wing and over-wing installations. 

These concepts were evaluated in FLOPS for their ability to meet mission requirements. Technology 

suites and airframe configurations were objectively evaluated on the SER basis to form a preferred 

package. 

In a final iteration with the NAS simulations, the selected N+3 concept (Figure 1-1), was modeled 

and re-evaluated in a sample Hartsfield-Jackson Atlanta International Airport (ATL) metroplex to identify 

the benefits of the system on a local level. The impacts of operating the vehicle at a reduced cruise speed 

compared to the reference vehicle, as well as the impacts of the fuel efficiency and mission performance, 

were evaluated on a system level to assess more realistic operational effects. 

Technology maturation plans for the preferred technology package were devised for the development 

of the technologies to maturity in a timeframe to support a 2030-2035 EIS. These roadmaps project the 

developmental requirements associated with maturing each technology to TRL 6 by the year 2025. It is 

estimated that this lead time will be required to implement these technologies into the design and 

development phase of the N+3 generation of commercial transport aircraft. 

NASA/CR—2010-216798 6
3 AIR VEHICLE DESIGN TOOLS AND PROCESSES 

3.1 Flight Optimization System (FLOPS) 

FLOPS is a tool developed by NASA Langley Research Center that allows for the preliminary design 

and analysis of flight vehicles [Ref. 1]. FLOPS combines nine separate modules that allow for the 

multidisciplinary design of an air vehicle. These modules include weights, aerodynamics, engine cycle, 

propulsion data scaling and interpolation, mission performance, takeoff and landing, noise footprint, and 

cost analysis, all controlled by an outer program control loop. 

The weights module uses empirical equations to predict the system and mission segment weights, and 

outputs a high-level weight summary. The module contains detailed inputs for a wide range of aircraft 

components and subsystems that allow for the calibration of their respective weights. These parameters 

can be further modified to reflect weight changes caused by the application of various technologies to the 

vehicle. Once properly calibrated using Northrop Grumman’s CONMAP (Section 3.4), a proprietary 

parametric weight estimation program, the FLOPS weights module outputs the weight summary. Center 

of gravity and balancing were also performed using the CONMAP program discussed in Section 3.4. 

The aerodynamics module stems from the Empirical Drag Estimation Technique (EDET), an 

algorithm based on user input which calculates drag polars for high-speed performance. The module 

contains input parameters used to calibrate the high-speed drag polars to achieve expected aerodynamic 

performance based on in-house databases and experience. Low-speed drag polars were input into this 

module for takeoff and landing, with corresponding reference values for geometric scaling as the aircraft 

is resized by FLOPS. 

Engine data tables provided, among other things, the thrust and fuel flow information required by 

FLOPS. The propulsion data scaling and interpolation module fills in any missing data in the engine 

tables, and uses both linear and nonlinear scaling to adjust them to the required sea level static thrust. 

While engine scaling is conducted automatically in this module, the user must ensure that the scaling does 

not exceed the limits imposed by the engine model. 

The mission performance module uses outputs from the weights, aerodynamics, and propulsion 

modules to compute performance, optimum climb profiles, optimum or user-specified cruise segments, 

and user-specified descent schedules. Other performance parameters may be specified or calculated in this 

module, including acceleration, turn, refueling, payload release, hold segments, and reserve fuel 

calculations. 

The takeoff and landing module computes balanced and landing field lengths, approach speed, second 

segment climb gradient, and missed approach climb gradient. The module generates detailed takeoff, 

climb, landing, and descent profiles that were subsequently used in calculating the aircraft noise 

characteristics. 

The final module contained in FLOPS pertinent to this study is the program control module. FLOPS 

can perform a point design based on a full set of geometric inputs, conduct a parametric study across an 

array of design variables, or optimize a geometry based on an input objective function. This optimization 

may use one of several algorithms, including Broyden-Fletcher-Goldfarb-Shano and Steepest Descent. 

The optimization routine in FLOPS was not used during later portions of this study due to its inability to 

converge over broad design spaces. The point design and parametric analyses were later used in 

conjunction with an internally-developed optimization tool to be discussed later in Section 3.2. 

3.1.1 Calibration Process 

In order to calibrate the FLOPS tool, an analysis was performed on the Boeing 737-800, for which 

weight and performance data was readily available. This process is detailed in Figure 3-1, and will be 

described in the following sections. As the design study matured and new vehicles began to deviate from 

NASA/CR—2010-216798 7
the initial design point, this process was repeated to ensure that the calibration remained valid. In this 

manner, confidence in the tool was maintained throughout the duration of the study. 

Figure 3-1. FLOPS Calibration Process With the 737-800 

3.1.1.1 Configuration Weight Calibration and Validation – The Boeing 737-800 was modeled 

in CATIA and FLOPS using information from the Boeing 737 Airport Planning Guide [Ref. 2]. Engine 

tables were obtained for the CFM56-7B27, the engine installed on the 737-800. Upon being modeled in 

CATIA, the 737-800 CATIA file was analyzed in CONMAP, producing a system- and component-level 

weight statement. This CONMAP statement was compared against publically-released weight 

information for this aircraft to verify its accuracy. In parallel to the CONMAP analysis, the configuration 

was modeled and analyzed in FLOPS, generating a similar weight statement. The weight statements from 

both tools were compared and segment level adjustments made in FLOPS to calibrate the component 

weights predicted by FLOPS to the higher-fidelity predictions of CONMAP to within a specified 

tolerance. As a result, the subsystem weights, overall vehicle empty weights, operating weights, zero fuel 

weights, and gross weights output by FLOPS were within one percent of those output by CONMAP. The 

weight statement comparison between CONMAP and the calibrated FLOPS tool is shown in Table 3-1. It 

must be noted that in the FLOPS analysis, the weights of the accessory gearbox, engine controls, and 

starting system are all included in the engine installation weight, and not individually broken down. 

FLOPS 737-800

CATIA File

FLOPS Single 

Point Analysis

737-800

CONMAP File

Initial 737-800

FLOPS File

737-800

FLOPS File

Weights Calibration & Validation

CATIA 3-D 

MODEL CONMAP

Aerodynamics Calibration

FLOPS Multi 

Point Analysis

737-800

FLOPS File

Publicly Released 

Performance Information

1

2

3

Aerodynamics Validation

Aerodynamic

Validation

Performance

Comparison

0.00

2.00

4.00

6.00

8.00

10.00

12.00

14.00

16.00

18.00

20.00

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

Mach

L/D or M *Max L/D

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

0.0000 0.0200 0.0400 0.0600 0.0800

CD

CL

0.0400 0.0600 0.0800 Historical CD

Aerodynamic Trends

Weight Statement 

Comparison

737-800

FLOPS File

Final 737-800

FLOPS File

Publicly Released B737-800 

Weight Information

IF FLOPS = CONMAP

IF FLOPS = PUBLIC

NASA/CR—2010-216798 8
Table 3-1. CONMAP and FLOPS Weight Calibration for the Boeing 737-800 

CONMAP FLOPS

 Weight (lb) Weight (lb) [lbs] [%]

Structure 51433.8 51743 309.2 0.60%

Wing 19230.6 19391 160.4 0.83%

Horizontal Tail 1411.2 1408 -3.2 -0.23%

Vertical Tail 1490.5 1487 -3.5 -0.23%

Fuselage 18055.2 18135 79.8 0.44%

Alighting & Arresting Gear 7289.4 7364 74.6 1.02%

Engine Section & Nacelle 3956.9 3958 1.1 0.03%

Miscellaneous 0.0

Propulsion 13259.1 13259.0 -0.1 0.00%

Engine Installation 10468.0 11246 -0.1 0.00%

Afterburner 0.0

Exhaust System 0.0

Accessory Gearboxes/Drives 562.7

Engine Controls 79.3

Starting System 136.0

Fuel System 1029.1 1029 -0.1 -0.01%

Miscellaneous (Thrust Reverser) 983.9 984 0.1 0.01%

Systems & Equipment 23883.0 24090.0 207.0 0.87%

Flight Controls 3421.3 3445 23.7 0.69%

Auxiliary Power Plant 934.0 933 -1.0 -0.11%

Instruments 656.6 660 3.4 0.52%

Hydraulics & Pneumatics 1418.4 1428 9.6 0.68%

Electrical System 2378.2 2400 21.8 0.92%

Avionics 1620.0 1634 14.0 0.86%

Furnishings & Equipment 12343.9 12491 147.1 1.19%

Air Conditioning & Anti-Icing 1010.7 999 -11.7 -1.15%

Load & Handling 100.0 100 0.0 0.00%

Weight Empty 88575.9 89092.0 516.1 0.58%

Crew and Baggage - Flight 420.0 420 0.0 0.00%

Crew and Baggage - Cabin 1050.0 1050 0.0 0.00%

Engine Oil 120.0 120 0.0 0.00%

Fuel, Unusable 138.0 138 0.0 0.00%

Passenger Service 1031.9 1040 8.1 0.78%

Clean Operational Weight 91335.8 91860.0 524.2 0.57%

Cargo - Containers 0.0 0

External Fuel Tanks & Provisions 0.0 0

Miscellaneous 0.0 0

Operational Weight 91335.8 91860.0 524.2 0.57%

Passengers 27540.0 27540 0.0 0.00%

Passenger Baggage 6480.0 6480 0.0 0.00%

Cargo - Containers (Baggage only) 1296.0 1307 11.0 0.85%

Cargo (Chosen to match TOGW) 1548.0 1548 0.0 0.00%

Fuel, Usable - Internal 46000.0 45917 -83.0 -0.18%

Takeoff Gross Weight 174199.8 174652.0 452.2 0.26%

Delta

NASA/CR—2010-216798 9
Validation of the FLOPS calibrations against CONMAP was needed to verify the calibration of the 

tools. The newly calibrated weight statement compared well to publically released weight data for the 

vehicle [Ref. 2]. Maximum ramp weights, takeoff and landing weights, zero fuel weights, empty weights, 

structural weights, and payload weights were within an acceptable tolerance, thus validating the weight 

calibration for additional tube-and-wing. A similar process was used to calibrate these tools for hybrid 

wing-body vehicles using a Northrop Grumman internal database of all-wing aircraft. 

3.1.1.2 Tube-and-Wing Aerodynamic 

Calibration – After calibrating and validating the 

vehicle weights, calibration of the FLOPS 

aerodynamics module followed. Takeoff and 

landing maximum lift coefficients and low-speed 

drag polars were set to historically accurate values 

for comparable tube-and-wing vehicles. The lowspeed drag polars were input into FLOPS such that 

they would be geometrically scaled. High-speed 

drag polars were calculated internally in FLOPS 

through the aerodynamics module, but were 

calibrated to match publically released performance 

characteristics for multiple payloads and ranges as 

shown in Figure 3-2 and Table 3-2 [Ref. 2]. 

Aerodynamic outputs from FLOPS were validated 

against historical trends for tube-and-wing aircraft 

of similar gross weight and passenger count. 

3.1.1.3 Hybrid Wing-Body Aerodynamic 

Calibration – The hybrid wing-body aerodynamic 

characteristics were calibrated against a proprietary 

database of Northrop Grumman all-wing vehicles. 

Low-speed drag polars and lift coefficients for 

takeoff and landing were input into FLOPS such that 

they would be geometrically scaled as the 

configuration was resized. High-speed drag polars 

were calculated internally in FLOPS, but were 

calibrated to the polars of similar all-wing 

vehicles. The drag divergence Mach number, 

maximum lift-to-drag ratio, and cruise lift 

coefficient were among other aerodynamic 

parameters calibrated in the FLOPS aerodynamic 

module against this internal database. 

3.1.2 Reference Vehicle 

For comparison and traceability during this 

study, a reference vehicle was configured as 

shown in Figure 3-3. This reference vehicle was 

derived from the Boeing 737-500 with the CFM56-3B1 engine. However, the aircraft was resized to meet 

the same requirements as the N+3 vehicle in terms of field length, passenger count, and range. The 

737-500 and CFM56-3B1 engine were modeled in FLOPS and were resized to meet the N+3 mission. In 

the resizing of the aircraft, technology level was held constant in addition to nondimensional 

characteristics including aspect ratio, taper ratio, and thickness-to-chord ratio. Engine cycle was held 

constant, and the engine was scaled in size and weight as necessary. The wing area and thrust increased 

over the 737-500, primarily driven by the 5,000-foot field length requirement. The resulting vehicle 

Figure 3-2. Boeing 737-800 Performance 

Calibration Points (Shown in Red) 

Table 3-2. 737-800 Performance Calibration Points 

NASA/CR—2010-216798 10
served as a reference from which the N+3 achievements 

were measured. Its characteristics are summarized in 

Table 3-3 and are compared against the 737-500 from 

which it was derived. 

Figure 3-3. Reference Vehicle 

3.2 Optimization Method 

Because the FLOPS’ optimizer has well-documented problems with converging over broad design 

ranges[Ref. 3, 4], an internally developed tool was used to optimize parametric FLOPS solutions on a 

user-specified objective function. For example, an objective function could be defined to minimize fuel 

weight only, or a weighted combination of empty weight and fuel weight. The optimization tool uses 

FLOPS’ parametric analysis mode to generate an array of vehicles with varying wing area, aspect ratio, 

sweep and thrust within specified boundaries. A parsing algorithm then sorts the resulting configurations, 

eliminating those that do not meet imposed requirements. For example, configurations that did not meet 

the balanced field length constraint were eliminated, as were designs with wingspans greater than 150 ft. 

The distilled configuration list is then output to an Excel file where the objective function is applied to 

each configuration to further sort the remaining vehicles and refine the boundaries of the next parametric 

analysis iteration. This process is repeated for several iterations, until the vehicle design tool converges 

within a user-specified tolerance on an optimum configuration. The objective function was chosen to 

minimize a combination of empty weight and fuel weight. Though fuel burn remains a primary target of 

the N+3 program, the decision was made to 

maintain some recognition that empty weight 

must be constrained. Without this constraint, 

FLOPS would create impractical 

configurations, most notably very large 

aspect ratio aircraft. Once an optimum 

configuration is reached for the given inputs 

(i.e., airframe configuration, engine, and 

technology application), it is then analyzed in 

FLOPS’ point-design mode to generate a 

detailed mission profile, including takeoff 

and landing trajectories, weight breakdown, 

and segment fuel reports. Figure 3-4 shows 

the multitude of solutions developed by the 

FLOPS parametric solver for four 

consecutive iterations. 

Table 3-3. Reference Vehicle 

Characteristics 

Figure 3-4. Optimization Methodology 

Units 737-500 Reference 

Vehicle

# Passengers [] 123 120

Range [nm] 2,400 1,600

Ramp Gross Weight [lb] 133,500 120,170

Empty Weight [lb] 68,860 67,350

Fuel Weight [lb] 42,186 25,048

Wing Reference Area [ft^2] 1,135 1,280

Wing Sweep [deg] 25 25

Wing Span [ft] 94.9 100.4

Wing AR [] 7.9 7.9

T/W Ratio [] 0.3 0.34

Max Wing Loading [psf] 117 94

Balanced Field Length [ft] 8,630 4,497

Landing Field Length [ft] 4,450 4,996

NASA/CR—2010-216798 11
3.3 Airspace Concepts Evaluation System (ACES) 

ACES is a tool that simulates nationwide air traffic management, flight, and airspace operations 

center functions [Ref. 5]. ACES simulates the air traffic system by using software that simulates hundreds 

of airports with Terminal Radar Approach Control (TRACONs), flight centers and sectors, thousands of 

flights gate-to-gate, and numerous airspace operations centers, as well as the interaction between these 

agents. ACES is able to simulate a realistic operating environment through complex multi-day flight 

schedules and flight plans, four dimensional (4-D) gridded winds, and adjustable operation conditions. 

Aircraft trajectories are realistically simulated through a four degree-of-freedom model that is based on 

laws of physics, aerodynamics, pilot-based control laws, and elliptic Earth trajectory propagation. ACES 

creates dynamic visualizations of the airspace and delays, and is equipped with extensive data recording 

and analysis tools. 

The tool allows the system-wide impacts of new aviation concepts to be analyzed, and can visualize 

how the NAS will handle future flight demands. ACES is able to analyze the effects of introducing N+3 

vehicles into current and future air traffic environments. It can also assess the exploitation of metroplex 

operations by integrating NextGen-enabled satellite airfields. The tool is used in this study to complete 

trades on cruise Mach number, range, and runway length for N+3 vehicle configurations and the resulting 

effects on the complex air space and operations environment. A diagram of the components and process 

used by ACES is shown in Figure 3-5. 

Figure 3-5. ACES Simulation Components and Data Flow Diagram 

3.4 Conceptual Mass Properties (CONMAP) 

CONMAP is a parametric feature-based weight estimation program developed using the Northrop 

Grumman aircraft database. This proprietary program provides high-level weight and longitudinal center 

of gravity estimates based on vehicle geometry, system definitions, and vehicle functions, as well as 

NASA/CR—2010-216798 12
allowing for the incorporation of special features not within the database. For initial calibration of the 

FLOPS tool, a CONMAP tool was used for calculating the mission segments: vehicle empty weight, 

operational empty weight, and basic mission takeoff weight. Factors within the FLOPS tool were adjusted 

for calibration on the initial starting points. Furthermore, CONMAP was used to calculate the center of 

gravity for wing placement on the preferred configuration. 

3.5 Numerical Propulsion System Simulation (NPSS) 

NPSS is a simulation tool licensed by Wolverine Ventures, Inc. that was originally developed at 

NASA Glenn Research Center for the analysis and design of complex systems, such as aircraft engines 

[Ref. 6]. As an object-oriented engineering tool, NPSS is easily integrated into multidisciplinary 

optimization tools and adaptable to many diverse applications and software. Existing software 

architecture allows for the simple modeling of aircraft engine components, including compressors, 

turbines, mixers, ducting, and combustors. Its flexibility allows for user-specified degrees of fidelity in 

the models of individual engine components. For example, if three dimensional (3-D) computational fluid 

dynamics (CFD) is available to predict the flow through a particular compressor, this software or its 

results can be interfaced with NPSS to serve as a performance model. At the same time, a simple 

polytropic efficiency can be used to calculate the turbine flow. This software has become an industry 

standard in engine development, and is the primary tool that was used to develop the N+3 advanced 

engine architectures. 

3.6 Acoustic Liner Design Process 

Design and analyses were performed of various inlet liners for N+3 vehicle propulsion systems as 

part of the noise reduction effort (Figure 3-6). Acoustic impedance models were utilized, validated 

through normal incidence and grazing flow impedance testing conducted at the Spirit Aerosystems flow 

duct experimental facility [Ref. 7]. Based on bypass fan propulsive sources predictions, these models 

calculated the impedance impact for various liner designs. Trade studies were performed to design a liner 

best capable of realizing the greatest insertion loss, thus reducing community noise at multiple flight 

conditions. These trade studies yielded an optimum conceptual acoustic liner based upon technologies 

projected to achieve sufficient TRL in the 2025 timeframe for the given propulsion system inlet nacelle, 

as reflected in the noise analyses of the N+3 vehicles. A summary of the major findings of this study can 

be found in Section 6.1.2. 

Figure 3-6. Acoustic Liner Conceptual Design Examples 

3.7 Model for Investigating the Detectability of Acoustic Signatures (MIDAS) 

MIDAS is a Northrop Grumman proprietary atmospheric acoustic detection analysis tool combining 

significant aspects of aircraft noise source generation, atmospheric propagation, and ground-based or 

airborne detection systems. It is a comprehensive model that for allows for the assessment of acoustic 

signatures and the establishment of design requirements and constraints. Given a flight profile, a series of 

algorithms are coordinated to generate a desired acoustic metric (e.g., dBA, DNL, PNL, EPNL, etc.). 

Noise source algorithms include exhaust mixing, broadband shock cell, combustion, turbomachinery, 

Facesheet

Septum

NASA/CR—2010-216798 13
trailing edge, discrete frequency rotor, propeller, and models to predict shielding and refraction, acoustic 

radiation and duct propagation, and acoustic attenuation and reflection. MIDAS generates noise source 

models, normalized far-field directivity maps, and integrated noise metrics for cumulative noise 

assessment as defined in FAA FAR Part 36. Some of the noise source generation modules have been 

adapted from NASA’s Aircraft Noise Prediction Program (ANOPP) program. 

The Advanced Sound Propagation in the Atmosphere program (ASOPRAT) is a hybrid ray trace/fullwave model prediction program that represents the state-of-the-art in long-range atmospheric sound 

propagation, and includes modules for atmospheric absorption, ground reflection, and multipath 

propagation [Ref. 8]. The software contains discrete frequency and one-third-octave band capability. 

Thorough testing and validation of the code was performed using data from a joint acoustic propagation 

experiment conducted at the White Sands Missile Range [Ref. 9]. 

A diagram of the process used by MIDAS to compute the required noise information is shown in 

Figure 3-7. Because of the level of detail involved in computing trajectory information at 0.5 second 

intervals required by MIDAS, acoustic calculations were reserved for the configurations with the most 

potential for noise reduction, and configurations which best represented the progression of noise reduction 

from the reference aircraft to the preferred configuration. 

Figure 3-7. MIDAS Flow 

3.8 System Effectiveness Rankings 

In order to ascertain the preferred configuration, system effectiveness ratings (SERs) needed to be 

defined, taking into consideration the 2030-2035 future scenario and the N+3 goals for each success 

metric. This section will define the SER considered for each of the N+3 metrics, as well as the overall 

SER for each configuration, based on the overall future scenario described in Section 4.1.2. 

SER definitions were necessary to quantify the relative successes and shortcomings of various 

configurations relative to each of the N+3 metrics. For the performance goal of achieving a 70% fuel burn 

reduction over the reference vehicle and mission, a relationship was defined between the achieved fuel 

burn, the reference vehicle fuel requirement, and the associated SER: 

NASA/CR—2010-216798 14
ref

ref

fuel MF

MFMF

SER

.0 70

M M

Eqn. 1

where MFref is the required mission fuel for the reference vehicle (25,048 lbs), and MF is that of the 

vehicle under consideration. SERfuel reaches one when the fuel burn for the configuration under 

consideration reaches 30% of that of the reference configuration. If the fuel burn exceeds that of the N+3 

goal, no extra credit is given, and the SER for this metric remains unity. However, penalties are assessed 

through negative values of SER if the fuel burn exceeds that of the reference configuration. 

Similarly, the emissions SER is defined by the following equation: 

6/

6/

.0 75 CAEP

CAEP

NOx NOx

NOx NOx SER N N

Eqn. 2

where NOxCAEP/6 is the maximum permissible NOx production based on engine OPR and rated SLS 

thrust, in grams per kN of SLS thrust at maximum power, as described in Section 6.2.1. Again, no extra 

credit is given for configurations which surpass the N+3 target of 75% reduction in NOx production 

compared to the CAEP/6 regulation. 

In order to quantify the SER for the balanced field length metric, the decision was made to use a 

simple step function which would give full credit to configurations which met the balanced field length 

requirement, and no extra credit to those that surpassed the target: 

BFL

SER floor BFL

5000

Eqn. 3

where ε is an arbitrarily small positive number. 

The final metric for which a system effectiveness rating definition is required is noise, and this SER 

was defined with respect to the Stage 4 noise requirement: 

Eqn. 4

With definitions in place for effectiveness ratings for each of the N+3 metrics, the mapping between 

the goals and the future scenario can be completed to define an overall system effectiveness rating. This 

definition is based on the relative importance of each of the N+3 goals in the framework of the overall 

future scenario. Using the weighting parameters derived in Section 4.1

 effectiveness rating for each configuration can be defined:

Eqn. 5

As defined, the total system effectiveness rating is most heavily weighted by fuel burn and emissions, and 

least affected by the balanced field length requirement, unless this takeoff capability is not met. 

71

EPNL 4 EPNL

SER Stage

EPNL

E E

SER SEREPNL SERfuel SERNOx 133SERBFL 0. 200 0. 333 0. 333 0.

NASA/CR—2010-216798 15

, the total system
4 FUTURE SCENARIO AND REQUIREMENTS DEVELOPMENT 

A future scenario analysis was performed in order to determine the overarching mission requirements 

for the N+3 vehicle. This analysis took into account mission type, vehicle size class, and the importance 

of the N+3 goals themselves in the 2030-2035 timeframe. The economic, political, and social 

environments, as well as available natural resource supply and demand, were investigated to estimate the 

scope of most-probable global scenarios in the timeframe of EIS. The fifteen scenarios identified in were 

reviewed and grouped into four categories based on their similar views of the probable future 

[Ref. 10,11,12,13]. An important guiding tenet was that no specific scenario would be adopted for 

primary use in the N+3 vehicle studies. This was done to produce a vehicle which, in the context of the 

N+3 goals, can also fulfill the needs of the future scenario regardless of the actual outcome. This 

approach should lead to a robust system that fills a broad, primary need in the NAS at the lowest possible 

risk for technology development funding. 

4.1 2035 Scenario 

The scenarios described in this section were developed by conducting research on existing studies, 

performing literature searches on major open issues, and holding technical interchange meetings to review 

findings and interpret the impact each scenario would have on the N+3 metrics. Three primary scenarios 

were developed across all of the research findings, in addition to a fourth scenario that was set aside due 

to its dire implications. Beyond illuminating the key aspects of the future of commercial transports, the 

primary outcome was the development of weighting factors intended to discern the value of concepts and 

technologies with respect to the N+3 goals. This drove the technology selection and initial ranking 

process to create suites applicable to a given airframe concept. A brief overview of the resources used to 

conduct this investigation is provided, along with descriptions of the three scenario groupings. Finally, the 

impact of the scenarios in terms of the N+3 goals is provided, and a quantitative estimation of an overall 

hedged scenario is presented. These scenario weighting factors were then applied to shape the vehicle 

requirements definition and technology selection process. 

4.1.1 Scenario Development and Key Drivers 

A total of fifteen existing scenarios were assessed to determine if any categorization existed for likely 

outcomes in the N+3 timeframe. These reports stem from several different organizations, chosen to 

encompass various interests in the products of scenario planning. Table 4-1 summarizes and organizes 

these resources by organization and scenario designations. 

The scenarios were then grouped according to key resource demand drivers and the general trends 

exhibited, based on individual central themes. The scenarios investigated had many secondary drivers in 

common, and while they were developed across a span of a decade, there were enough similarities across 

the groupings to warrant categorization. As shown in Figure 4-1, the four categories relating to future 

resource demands are: (a) reducing demand, (b) balancing supply/demand, (c) increasing resources, and 

(d) collapse in demand. 

NASA/CR—2010-216798 16
Table 4-1. Specific Scenarios Used to Develop the N+3 

U.S. Airspace Overall Scenario [Ref. 10, 11, 12, 13] 

Figure 4-1. Future Scenario Categorization Quadrants 

This grouping provided a foundation with which to evaluate vehicle performance against a smaller 

subset of key market drivers and external constraints, including environmental factors such as noise, 

pollution, and global CO2 levels. The resource demand projections in Figure 4-2 reveal the nature of the 

Organization Report/Scenario Title 

A) JPDO Futures Working Group “Futures Working Group Final Report (Draft)”, 2004 

1. Is it hot in here or what? 

2. Storm Clouds 

3. Markets Rule 

4. Asia’s Century

5. Terror Uncontained 

B) National Research Council (NRC) “Maintaining U.S. Leadership in Aeronautics: ScenarioBased Strategic Planning for NASA's Aeronautics 

Enterprise”, 1997

1. Pushing the Envelope

2. Grounded

3. Regional Tensions

4. Trading Places

5. Environmentally Challenged

C) World Business Council for 

Sustainable Development (WBCSD) 

“Exploring Sustainable Development”, 1997

1. FROG! First Raise Our Growth 

2. Jazz – Dynamic Reciprocity 

3. GEOpolity – Sustainable Guidance to Market 

D) Shell “Shell Energy Scenarios to 2050”, 2008

1. Scramble! 

2. Blueprints 

NASA/CR—2010-216798 17
four scenario sets. Since the collapse scenario had little effect on vehicle trade studies (i.e., no technology 

set could achieve a meaningful solution), it was not considered further. 

Figure 4-2. Simplified Future Resource Trends for Various Scenario Quadrants 

Estimating the future of energy and resource usage involves the synthesis of many factors affecting 

macroeconomic productivity. Political instability and opposition to collaboration will affect trade 

balances, tariffs, resource markets, currency markets, and budget allocations. Extreme forms of political 

change include the U.S. being excluded or self-isolated from major developing and emerging markets. 

World markets, independent of political interdiction, have a profound effect on traffic demands and the 

availability of resources at any given time, as shown by recent oil price fluctuations. The coupling of 

other highly-fluctuating energy resources, for example natural gas, have an unknown effect on oil 

availability and prices, especially while new methods for natural gas extraction and processing are being 

developed and used. Furthermore, military acts and conflicts can produce short- and long-term effects, 

ranging from minor regional resource shortages to essentially indefinite global effects. Technological 

advances in areas other than aviation toward sustainable or renewable energy sources could greatly 

contribute to resource availability. Since any of these drivers could have a large impact, and their 

certainty of occurrence is very difficult to predict, it is important to not delve into the possibility of any 

one particular scenario, but to investigate the macroeconomic impacts across the range of possibilities. A 

robust vehicle design, then, is one that will perform as well as possible in terms of the N+3 goals, under 

any of the likely outcomes. Consideration was then given to the macroeconomic effects of primary 

concern to commercial air vehicle traffic in the NAS and with regard to the N+3 goals: 

Availability of crude oil (impact on N+3 fuel burn reduction goal) 

Sensitivity of communities and populations to aircraft noise and emissions (impact on N+3 noise, 

LTO NOx goal) 

Airfield resources required to sustain or improve NAS system performance (impact on N+3 field 

length capability). 

4.1.1.1 Resource Availability – Over the past decade, the U.S. has experienced unprecedented 

influences on its aviation market, including terrorism, projections of doubling or tripling air travel 

demands, significant swings in energy costs, growing environmental issues, ATM saturation, a global 

financial crisis, “buy American” sentiments, and the continuing development of a national plan for air 

transport. These variables become even more important when recognizing that the aerospace sector has 

provided substantial benefits to the national economy. Aerospace technology is consistently the fourth 

strongest export item when ranked by total export dollars ($76B in 2007), and represents 6-8% of total 

exports [Ref. 14]. In 2008, civilian aircraft engines were the fifth largest export growth category (up 

$2.5B from 2007). 

NASA/CR—2010-216798 18
The future of air transportation is likely to be dominated by national strategies and objectives in 

energy, the environment, the economy, and defense. Oil and other sources of energy obviously play 

important roles in the world – and in aviation in particular. Figure 4-3 shows the International Energy 

Agency (IEA) assessment of historical energy consumption and projections out to 2030. World energy 

demand expands by 45% between now and 2030, an average rate of increase of 1.6% per year. Today’s 

world oil consumption is just under 80 million barrels per day (Mbpd), and consumption is projected to 

grow through 2030. 

Figure 4-3. Projected Global Energy Needs Through 2030 [Ref. 15] 

In considering sources of oil (Figure 4-4), unconventional sources are a growing fraction of the 

supply. After 2010, by far the greatest sources of oil are those which have yet to be developed [Ref. 15]. 

A total of 64 Mbpd of gross capacity needs to be installed between 2007 and 2030, a number that is six 

times the current capacity of Saudi Arabia, in order to meet demand growth and offset the decline of 

existing oil fields. The primary message of the IEA is that the continued growth of energy consumption is 

not sustainable. 

Figure 4-4. Production Estimates of Current and Future Crude Oil Sources [Ref. 15] 

The U.S. is the largest consumer of oil, at over 20 Mbpd [Ref. 16]. This is more than twice the second 

largest consumer, China. U.S. oil production peaked in the 1970s and 1980s, and declining production 

with increased demand is addressed by oil imports. Today, roughly half of the oil consumed in the U.S. is 

imported from overseas, and this fraction continues to grow. Such dependencies make the U.S. and the 

aviation industry susceptible to shortages and attendant fuel price increases regardless of world supply 

levels. Entities that control the oil supply can greatly influence the capability of the U.S. airspace to meet 

growing traffic demand. 

NASA/CR—2010-216798 19
4.1.1.2 Scenario Descriptions – After carefully studying the fifteen scenarios, they were grouped 

based on which of the four future energy projections (Figure 4-2) each scenario supported. The scenarios 

generally aligned with these four projections. As shown in Table 4-1, only the National Research Council 

(NRC) and Joint Planning and Development Office (JPDO) considered the Collapse scenario. Neither of 

the studies outside of the air transport sector considered this to be a reasonable outcome. Given that 

collapse implies an almost complete breakdown of air transport, this scenario was discarded from further 

consideration in this study. There would appear to be no reasonable mix of technologies that are relevant 

to this projection. 

Another simplification was based on the similarities between the 1997 NRC scenario study and the 

2004 JPDO Futures Working Group scenarios. Effectively, since the JPDO scenarios represent a more 

recent, refined view of the NRC scenarios, the latter was dropped from further consideration. The 

following sections detail the three final scenarios, present supporting data from other studies, and discuss 

their implications in brief. 

1. Scenario 1: Bright Bold Tomorrow (BBT) 

In the framework of this scenario, new approaches to energy initiated in the early 21st century now 

pervade the transportation system. As a result, reasonably-priced carbon-based biofuels from algae farms 

are now widely available. The world economy has more than recovered from the downturn of the first 

decade, and renewed prosperity places high demand on global travel. New investment in airport 

infrastructure has saturated the hub-and-spoke system, and widespread point-to-point transportation is 

now available. Smaller, single-pilot aircraft dominate the transportation system, and the only realistic 

solution is to leverage smaller local airports and closer spacing for aircraft using larger airports. 

Buyers want to be able to purchase whatever they want, wherever they are and whenever it is 

convenient for them. Companies that were not able to respond quickly to rapidly changing demand have 

disappeared. This frenetic pace has helped to feed a strong economy in the U.S. and the rule of markets 

has spread increasingly to Asia and Europe. The environment, already deteriorating in many places by the 

turn of the century, is now approaching crisis levels in Asia. 

In a 2006 publication, ExxonMobil paints a rosy future of business-as-usual which supports a 

scenario of ever-expanding air transport [Ref. 17]. By 2050 they see daily worldwide oil production at 

110 million barrels per day oil equivalent (MbpdOE), which is well beyond production predicted in all the 

other scenarios (Figure 4-5). Their projection is, on average, a 1.5% yearly growth rate for all fossil fuels, 

as well as similar growth in other sources of energy. 

Figure 4-5. ExxonMobil 2006 Projections of Fossil Fuel Growth [Ref. 17] 

Attendant with their projections of energy growth are worldwide emissions of CO2 that increase 

dramatically with application of 1990-era technology, almost doubling over a 50-year period from 1980 

NASA/CR—2010-216798 20
to 2030 (Figure 4-6). They project three 

solution paths which are consistent with the 

growth scenario: focus on CO2 sequestration 

technologies, reduce the energy intensity of 

the world’s economies, and invest in bringing 

alternative sources of energy online at a more 

rapid pace. To counter the growth in global 

CO2 projected to continue through 2030, an 

aggressive application of existing 

technologies, preceded by an aggressive 

development of new technologies, must be 

pursued. For the technology set to be available 

by the N+3 timeframe, the development of 

multiple technology paths that will hedge 

against technical, physical, and political 

hurdles must begin as soon as possible to 

reduce development costs, and ensure successful deployment. A byproduct of this technology application 

is the implementation of new technologies that reduce (across all fuel-consuming industries) the specific 

energy required per dollar of gross domestic product. New manufacturing incentives and the focus on 

nonenergy-intensive production techniques would raise standards of living while consuming hydrocarbon 

fuels at a reduced rate. This would presumably keep fuel prices low and spawn increased growth in air 

transport, driving air traffic levels to the upper end of projections. 

2. Scenario 2: Not In My Backyard (NiMBY) 

A combination of new energy sources and new approaches to energy conservation has brought supply 

and demand into a long-term balance. The energy projections for NiMBY (Figure 4-7) are more sanguine 

than those in the previous scenario due to early attention to the peak-oil challenges. By reducing 

consumption early on, peak oil is stretched across 

three decades through 2040, allowing time for 

alternative sources to come online. Air 

transportation is growing slowly, but remains a 

strong component of the overall transportation 

mix. The push to greater efficiency and energy 

conservation has reinvigorated the cities, leading 

to much higher population densities. Land for 

airports is at a premium, and the luxury of large 

buffer zones between residential areas and 

airports is not viable. Local populations are 

becoming increasingly intolerant to noise and airport emissions such as NOx, soot, particulates, and 

hydrocarbon emissions. Strict rules evolve first in the European Union and California, and the bar for 

noise and local emissions regulation is set very high. 

Diverse players form alliances to solve social and environmental problems in the most pragmatic way 

possible. This is a world of social and technological innovations, experimentation, rapid adaptation, 

voluntary interconnectedness, and a powerful and ever changing global market. Many stakeholders are 

involved, encouraged to step onto the economic stage by advances in information technology. That stage 

is characterized by a global free market, sound legal systems, and a respect for property rights. The public 

is made aware of transgressions and quickly acts against companies or countries that violate standards. As 

a result, effective, market-driven, demand-side efficiency measures emerge more quickly, and CO2 

management practices spread. Carbon trading markets become more efficient, and CO2 prices strengthen 

early. Energy efficiency improvements and the emergence of mass-market electric vehicles are 

Figure 4-6. ExxonMobil 2006 Projections of 

Near-Term CO2 Emissions [Ref. 17] 

Figure 4-7. Projections of Energy Sources 

for NiMBY Scenario [Ref. 13] 

NASA/CR—2010-216798 21
accelerated. The rate of growth of atmospheric CO2 is constrained, leading to a more sustainable 

environmental pathway. 

In the NiMBY scenario, it is recognized that air travel serves the entire population and must be 

sustained for our standard of living to persist. It is clear that air travel is vital to all elements of the 

population [Ref. 18]. The most difficult challenge in this scenario will be limiting the impact of air travel 

near major airports or airports in densely populated areas to acceptable limits. Given the need for air 

transportation, the development of environmentally-conscious communities, and the adaptation of more 

energy-conscious lifestyles, the key question becomes one of airport location. In its report to the United 

States Congress, the FAA/Partnership for Air Transportation Noise and Emissions Reduction 

(PARTNER) coalition states [Ref. 19]: 

“In 2025, significant health and welfare impacts of aviation community noise and local air quality 

emissions will be reduced in absolute terms, notwithstanding the anticipated growth in aviation. 

Uncertainties regarding both the contribution of aviation to climate change, and the impacts of aviation 

particulate matter and hazardous air pollutants, will be reduced to levels that enable appropriate 

action.”

Local impact is the dominant factor in expansion of air transport. Five million people live within 

55dB DNL areas, and noise is the single biggest local objection to airport expansion [Ref. 19]. Emissions 

of nitrogen oxides, carbon monoxide, unburned hydrocarbons, and particulate matter from a variety of 

airport sources contribute to local air quality deterioration. Aviation operations below 3,000 feet 

contribute 0.4% to the total national NOx inventory. For unburned hydrocarbons and particulate matter, 

airport emissions have the potential to be a public health issue rather than simply a local nuisance. In this 

scenario, the aviation community is driven to take action by a proliferation of local rules and legal barriers 

to growth. 

3. Scenario 3: King Carbon 

The confluence of peak oil and global warming 

concerns results in the imposition of strict “carbon 

taxes” to limit the burning of fossil fuels (Figure 4-8). 

Years of complacency and a business-as-usual 

approach has led to the realization that something 

must be done quickly. Insufficient investment in 

alternative fuels and sequestration technologies leaves 

little time to invest in alternative fuel infrastructure. 

Ground transport and electricity generation at least 

have some alternatives; however, air transport must 

address skyrocketing fuel costs with new approaches 

to conservation and efficiency. Global alliances 

cutting across developed and undeveloped countries 

set global standards for carbon-based fuel 

consumption. 

Long-held fears of global warming effects surface around the globe, though there still remains 

rampant disagreement on the causes. In spite of the very real and present effects of global warming, the 

world continued for years to respond with a slow and often low-magnitude reaction. The environment has 

deteriorated to the point that an international organization has been formed, charged with aggressively 

addressing global environmental conditions, and with commensurate authority to coerce cooperation from 

nations. By 2050, there is evidence that the darkest predictions about global warming are actually nearer 

to the truth than the more optimistic ones. 

The impact on energy resources becomes increasingly acute and unpredictably swift. Despite 

increasing rhetoric, action to address climate change and encourage energy efficiency is pushed into the 

future, leading to largely sequential attention to supply, demand and climate stresses. Demand-side policy 

Figure 4-8. Historical Global Crude Oil 

Availability and Production [Ref. 20] 

Accumulated Oil 

Discoveries

Production

NASA/CR—2010-216798 22
is not pursued meaningfully until supply limitations are acute. Likewise, environmental policy is not 

seriously addressed until major climate events stimulate political responses. Events drive late, but severe, 

responses to emerging pressures that result in energy price spikes and volatility. 

The overriding impact of carbon-based fuels in this scenario are the compounding impacts of 

dependence and the consequences of dependence. On one hand is the projection of peak oil, with no way 

to back down gracefully from the consumption peak unless aggressive preliminary action is taken. On the 

other hand are concerns about carbon dioxide emissions and their impact on global climate. Alone, either 

is a serious issue. Together, they increase the probability of this scenario becoming reality. The 

fundamental issue is that although oil discovery rates peaked in 1964, demand has continued to grow 

[Ref. 20]. Figure 4-8 compares accumulated oil discoveries with production through about 2004, with 

cumulative production beginning to approach cumulative discoveries. Alternative sources of oil exist, but 

the cost of making them a viable option begins to approach the value of the commodity. 

The second part of the impact of carbon-based fuels beyond market factors is the influence of CO2

production on global climate. In a recent paper by Marais et al [Ref. 21], the long-term impact of a step 

change in atmospheric emissions was examined (Figure 4-9). In this study, a number of emission 

constituents were assumed to be injected into the atmosphere over a short period of time. The impulse 

response in terms of unit global average temperature change and relative effect on GDP were estimated. 

On the left can be seen the temperature response, which can be divided into two groups. The majority of 

emission constituents have an initially large impact, but that impact decays to a negligible level within 

fifty years. This includes NOx, water vapor and cirrus clouds, and various sulfur compounds. After fifty 

years, the total emission footprint is dominated by CO2, and the decay rate of this constituent is very 

small. On the right, the cumulative impact on gross domestic product (GDP) is computed, and the 

transient of carbon emissions can last beyond 300 years. 

Figure 4-9. Impact of Impulse-Response Input of Various Greenhouse Gases 

on Atmospheric Temperature [Ref. 21] 

4.1.2 Scenario Synthesis and Impact to Vehicle Optimization 

The primary use of these detailed scenario descriptions was to illuminate the impact the future may 

hold on commercial air vehicles, NAS traffic, and the N+3 goals, while ignoring the drivers which could 

stem from a number of sources not considered, including the development of personal air vehicles, 

undiscovered power sources, etc. For instance, if personal air vehicles become pervasive before the N+3 

timeframe, then the major impact would be a reduction in NAS traffic. Likely, the impact would be small, 

since ranges and speeds would be very limited, but nevertheless, the effect of this particular example is 

well within the limits of the traffic growth levels predicted by the three scenario models. The scenarios 

strive to propose a manageable worst-case future set that covers the most dire decreases in air traffic 

demand due to huge price increases, as well as the most challenging economic growth levels that drive 

huge demand increases. 

NASA/CR—2010-216798 23
These three scenarios broadly address the spectrum of 

alternative futures and map directly to future fossil fuel 

projections. NiMBY is tied to airport site location near 

populated areas, and thus affects NOx emissions and 

community noise. King Carbon is primarily a fuelconstrained scenario, and is fundamentally tied to fuel burn 

and global emissions. The Bright, Bold Tomorrow scenario 

projects a large expansion in air travel demand, and is thus 

most closely tied to the field length goal, with little or no 

linkage to noise, LTO emissions, or fuel burn. In an attempt 

to make this mapping a quantitative, traceable effect, the 

impact of these scenarios to each of the N+3 goals was 

assigned on a percentage basis, with 100% allocated among 

the four goals. Figure 4-10 indicates these distributions as weighting factors for each of the three 

scenarios, which have specific implications that will influence the selection of vehicle technologies and 

mission requirements. 

NiMBY requires that a viable aircraft meet local emissions requirements that may vary between 

airports, and strict noise requirements as housing approaches airport boundaries. The focus of these 

vehicles would be low-noise, low-emissions operations. It would be critical that airlines bring fleets of 

new-technology aircraft online in a timely fashion to meet the new local requirements. 

King Carbon is predicated on the high cost of fuel, driven by shortages and taxes. A viable solution 

would have a diverse fuel mix, including synthetic and bio-derived fuels, as well as lighter, more efficient 

aircraft. Airspace solutions must be integrated with vehicles, as efficient flight paths would be mandated. 

As with NiMBY, airlines must be able to bring fleets of new-technology aircraft online in a timely 

manner to meet the new efficiency requirements. 

The Bright, Bold Tomorrow scenario follows a more business-as-usual trend, with the ability to use a 

wide range of airports. This scenario implies complex airspace management concerns, and emissions 

issues driven by a rapidly growing air travel industry. 

In support of the goal of this study, which is to define a single vehicle system that best achieves the 

N+3 metrics in the context of the 2030-2035 scenario, a hedged approach yielded a set of relative goal 

weightings. These factors were applied to initial technology selections through the technology 

effectiveness metric, TER, and the vehicle optimization metric, SER, to provide a vehicle and technology 

package best suited to a broad set of likely requirements. Figure 4-10 shows the values used to initiate the 

technology QFD process to select applicable best-suited technology suites, as well as begin to define 

vehicle requirements based on the traffic demand levels predicted by each scenario. This relative 

weighting of the scenarios reflects an average of the three scenarios previously described, and can be seen 

to more closely resemble the King Carbon scenario than the others. This is due to the relative importance 

of the emissions and fuel burn metric across all scenarios. 

4.2 Initial Requirements Development 

Of the requirements necessary to initiate a design optimization, including field length capability, 

vehicle range, cruise speed, and passenger count (payload), it was necessary to address the takeoff and 

landing field length capability before the others. This is a consequence of the metroplex concept being a 

tradable N+3 goal in the study. Since the field length requirement is dependent on (1) the traffic levels 

projected by the future scenario and (2) future runway resources enabled by NextGen, a study of these 

two aspects of the NAS were performed first to provide the basis for investigating field length. The goal 

was to estimate performance of the NAS in terms of the ultimate mission of delivering the future traffic 

demand levels predicted in the scenario study. The question of adopting metroplex operations (and to 

what extent) was assessed by comparing changes in aircraft delay at various traffic levels in two different 

Figure 4-10. Synthesized Scenario

Impact on N+3 Goals 

NASA/CR—2010-216798 24
cases. One used today’s baseline airport infrastructure, while a second incorporated metroplex resources 

enabled by NextGen. These simulations were performed with an N+3 surrogate vehicle to estimate the 

likely vehicle characteristics for that timeframe. 

Assessment of the cruise speed requirement was made using simulations that decreased cruise speed 

and showed the effects on delay. Passenger sizing was determined by investigating the flight distributions 

currently in the NAS, and reconciling field length requirements with the average mission type (range) of 

these vehicles to make use of the metroplex concept. This was found to be necessary to serve the 

challenging traffic demand levels predicted for the N+3 timeframe. The simulation approach and 

assumptions used to construct these tradeoffs is described prior to discussing results. 

4.2.1 Methodology 

The methodology used to analyze the advanced N+3 vehicle in the NextGen-enabled infrastructure 

had the goal of providing a system-wide assessment of passenger throughput and delay using the Airspace 

Concept Evaluation System (ACES), described in Section 3.3. Figure 4-11 shows the process that was 

used to assess system-wide performance. AvDemand [Ref. 28] was used to create future schedules using 

estimates of future passenger demand [Ref. 22]. A description of AvDemand use in creating flight 

schedules for the N+3 timeframe is described in Subsection 4.2.1.1. A tail-tracking algorithm was then 

used to link together the flights in the AvDemand-created schedules into airframe-based itineraries. The 

Base of Aircraft Data (BADA) contains specific aircraft performance data used by the ACES high-fidelity 

trajectory generator. Finally, the airspace and airport capacities were designed to include NextGen 

assumptions. 

Figure 4-11. Simulation and Modeling Approach for This Analysis 

A representative vehicle was chosen for analysis of N+3 vehicle performance and its impact on the 

National Airspace System. The targeted markets consisted of origination and destination (O/D) pairs 

serviced by the Boeing 737 and other similar aircraft. The Boeing 737 was selected based on similarities 

in aircraft size and target market. Separate future schedules were created and simulated for the reference 

vehicle, similar vehicles, and the advanced N+3 vehicle and the results compared. The differences were 

used to assess the impact of the advanced vehicle configurations on the NextGen-enabled NAS. 

Performance metrics included fuel efficiency, thrust, cruise speed, and cruise altitude performance 

profiles. System-wide comparisons included passenger throughput and delay. 

A bypass alternative was used to assess the effect of balanced field length (BFL) on passenger 

throughput and delay. The bypass alternative seeks to balance the demand to capacity by using policybased incentives and regulatory measures to shift demand to regional, underutilized facilities [Ref. 23]. 

Use of these existing metroplex assets was assumed via NextGen improvements such as Virtual Towers 

which would be implemented in the non-towered airports and allow reasonable hourly departure and 

arrival rates, estimated for this study at 40/hr. 

AvDemand ACES System Performance

Metrics

Tail Tracking

Algorithm

Future

Schedule

Seed day

Thursday, July 13, 2006

Airspace

Capacities

Airport

Capacities

NextGen Assumptions

BAse of Aircraft

Data (BADA)

NASA/CR—2010-216798 25
4.2.1.1 Baseline and N+3 Demand Set Generation – To generate the future demand sets, a 

representative “seed” day was chosen from which to extrapolate traffic patterns. Thursday, July 13, 2006, 

was chosen because summer is a busy season for air travel and Thursdays are a busy day of the week. 

Also, July 13 is not associated with a holiday weekend and can therefore be reasonably assumed to be 

representative of a busy travel day in the summer of 2006. AvDemand was used to grow the seed-day 

schedule using the expected passenger growth rates that are published annually in the FAA’s Terminal 

Area Forecast [Ref. 24]. These growth rates extend 15 to 20 years into the future and are based on factors 

such as expected growth rates for population and income. 

Expected passenger counts are converted into scheduled flights between city-pairs by using what is 

known as a Frataring algorithm. When additional flights are needed to handle extra enplanements in the 

future, AvDemand duplicates an existing flight in the seed day, slightly offsetting its departure and arrival 

times from those of the original flight. This process replaces individual flights in the seed day with 

multiple flights in the future demand set. The added flights are only offset from each other by a minute or 

so in order to replicate the operators existing business model. For instance, there are often many flights 

between certain city pairs early in the morning and late in the afternoon, but not in the middle of the day, 

as that is when airlines believe passengers want to fly. AvDemand preserves this business model; 

however, it also means that the forecasted demand sets probably contain slightly more delay than might 

actually be realized in the future, as the operators might change schedules to avoid congested peaks of 

arrivals or departures. Due to this approach, future demand sets distributions are similar to today’s, where 

traffic is concentrated at 30-40 major hubs. 

4.2.1.2 Airframe Substitutions – The seed day aircraft distribution was determined through 

AvDemand and used to develop the baseline and N+3 advanced vehicle usage rates. AvDemand’s 

passenger-weighted demand generation was used, which first converts flights to passengers and then 

redistributes passengers using a Gaussian distribution. The demand that is created is based on the fleet 

mix (percentage of market share for heavy, medium, large, and small) and aircraft type. The aircraft usage 

rates were adjusted to represent a hypothetical share by N+3 aircraft in the future flight schedules around 

the 2040 timeframe. The changes took into account: (1) the expected production rate of the N+3 aircraft 

(~4,000 available then), and (2) the retirement of present-day aircraft types. The JPDO fleet forecast was 

used for guidance during the development of the schedules. Examples of aircraft in the 2006 schedule 

assumed to be phased out in varying degrees are the MD-87, DC-9 and its variants, as well as variants of 

the B737. 

4.2.1.3 Tail Tracking – A tail tracking algorithm was used to connect multiple flights in the N+3 

and baseline demand sets. Tail tracking the demand sets was required to ensure adherence to aircraft 

turnaround times and to account for propagated delay in the system. By accounting for propagated delay, 

a tail-connected itinerary gives a more realistic estimate of delay when run through a simulation like 

ACES. In the current NAS, propagated delay accounts for about 30% of all delay [Ref. 25]. 

In a demand set that is not tail-tracked, every aircraft is at the airport and ready to take off at its 

scheduled departure time; it is only delayed because of congestion at the departure airport, destination, or 

in the airspace it must fly through to reach its destination. In reality, however, aircraft travel on itineraries 

between multiple city pairs and flights often take off late because the aircraft was delayed at some 

bottleneck earlier in the day. For example, an aircraft that departs an hour late from one airport will arrive 

an hour or so late at its destination. Depending on how much slack is in an airline’s schedule, this aircraft 

will likely depart late to its next destination, and so on throughout the day. 

The tail tracking algorithm generates schedules with a median of five flights per day. Figure 4-12 

shows a comparison between the tail tracking algorithm used to generate the N+3 and reference demand 

sets and the actual schedule for Continental Airlines derived from Bureau of Transportation data. 

4.2.1.4 NextGen Assumptions for N+3 Analysis – The airport and sector capacities used to 

simulate the N+3 vehicle in NextGen were provided by the JPDO, and assume NextGen to be fully 

implemented by 2025, with only small capacity increases after that year. These capacities include all 

known new and planned runway construction, as well as all JPDO operational improvements. Airport 

NASA/CR—2010-216798 26
capacities under NextGen have been modeled in detail by the JPDO, with NextGen operational 

improvements increasing average capacity for the top 35 airports by approximately 45%, although there is 

considerable variation around that average for the individual airports. Similarly, en route sector capacities 

increased by a factor of 1.7 when all NextGen en route operational improvements were modeled 

[Ref. 29]. 

4.2.2 Traffic Growth Implications on 

Metroplex Operations and Field Length 

Requirement 

The scenario task revealed the 

importance of determining vehicle 

requirements that affect the airspace 

efficiency so that large potential 

increases in air travel can be 

accomplished. The availability of oil and 

economic markets will mainly drive the 

fuel and operational efficiencies needed 

to maintain a viable fleet. Figure 4-13 

shows steady and substantial progress in 

increasing fleet efficiency over the past 

fifty years. The primary contributor is 

engine efficiency, driven by increasing 

bypass ratio. The inclusion of the last of 

the piston-engine propeller transports 

shows that the past fifty years have been spent returning to the fuel efficiencies of the propeller era. One 

must ask whether a lower Mach short haul transport aircraft is ultimately a better answer. The answer 

depends on the future traffic levels predicted, and whether or not the metroplex operations model is 

needed to meet this traffic demand, beyond the improvement assumed by NextGen with today’s 

infrastructure alone. Incorporating metroplex resource alone enabled by NextGen will increase throughput 

by fuller resource utilization. 

More than half the gains in fuel efficiency over the past fifty years have been related to propulsion 

systems. The remaining gains have come from higher structural efficiencies and reductions in 

aerodynamic drag, areas which still hold promise for even further gains. Most of the drive has been to 

make turbofans more efficient, and drive toward larger bypass ratios, but these are still are far less fuel 

efficient than equivalent turboprops or piston-props. Attendant cruise speeds are much lower for these 

aircraft though, so the efficiency benefits must be weighed against the designated mission for this vehicle: 

to best meet the future NAS operational efficiency requirements. 

Figure 4-13. Commercial Aircraft Fuel Efficiency Historical Data Indicating Piston-Prop 

Performance [Ref. 26] 

Piston-props

*MJ/ask = Mega-Joules Per Available Seat-Kilometers

Data from Lee (2001)

Data from IPCC (1999)

Micro analysis

A380

Figure 4-12. Comparison of Daily Flights per Airframe 

NASA/CR—2010-216798 27
4.2.2.1 Scenario-Based Traffic (Enplanement) Predictions – Based on the collection of 

scenario predictions noted above, air traffic growth levels were 

predicted for each of the scenarios. Table 4-2 indicates the traffic 

growth levels associated with each of the three individual 

scenarios relative to a 2008 air traffic baseline. These values are 

based on our own interpretation of the possible range of 

economic drivers on traffic growth. However, reports over the 

last decade have issued predictions across this range from 2-3 

times growth levels, for even earlier time periods [Ref. 27]. 

To compare this to historical data, the enplanements in the U.S. since the jet era begin are shown in 

Figure 4-14, plotting a quadratic extrapolation of the historical data against both the JPDO projection of 

approximately 1.8 times increase for 2040, with the BBT scenario projection of 4% per year growth. 

Clearly, the historical projection is increasing at nearly quadratic rates and, as long as technology and 

infrastructure maintain current proportions, would likely continue this growth rate at a minimum. The 

BBT air traffic growth of approximately 2.7 times by 2040 provides the most challenging environment 

with which to provide adequate infrastructure capacity and compatible air vehicle requirements, and is in 

line with the range of other recent predictions shown in Figure 4-14(b) [Ref. 27]. 

Figure 4-14. (a) Historical Passenger Enplanements Since 1952; (b) Future Traffic Predictions 

Comparison Ranges [Ref. 27] 

This high level of growth of the BBT scenario of 2.7 times 2008 levels is that adopted for the upper 

bound of the NAS simulations performed in this study is in line with the philosophy that an N+3 vehicle 

should be able to meet the most strict requirements as predicted by the future scenarios, and as demanded 

by the N+3 performance metrics. The baseline 1.8 times JPDO demand levels were also investigated for 

comparison. 

4.2.2.2 NextGen Infrastructure Assessment – The impact of the requirement to meet high future 

demand levels on metroplex needs was assessed by investigating the effects of traffic levels on future 

delay levels. A baseline comparison without implementing metroplex assets, but including NextGen 

improvements was made. NextGen represents a transformation of the air traffic control system from its 

present highly manual, analog, ground-centric state into one in which digital technology, information 

sharing, and automation are ubiquitous. This transformation affects the gamut of aviation activity, from 

air transport service business-models through airspace rulemaking [Ref. 28]. The relevant impact on this 

study is the increase in sector air capacities in FAA’s Monitor Alert Perimeter (MAP) of 1.7 times, and 

visual flight rules (VFR) departure and arrival capacity increases of about 45%. 

The effect of these increases alone due to NextGen implementation are shown in Figure 4-15, 

compared to a fleet with an assumed N+3 “like” surrogate vehicle. This was done to estimate the 

sensitivity without an actual N+3 preferred configuration at this point in the design process. It was 

modeled based on variations from an N+2 vehicle model and included a shorter range capability to take 

(a) (b)

Table 4-2. U.S. Air Traffic Growth 

Rates by Future Scenario 

Scenario %/yr 2035

NIMBY 2% 1.6x

KC -2% 0.6x

BBT 4% 2.7x 

NASA/CR—2010-216798 28
advantage of metroplex operations by using airports within 70 nm of an Operational Evolution 

Partnership (OEP) hub, and with runways at least 100 feet wide and 5,000 feet long. The N+3 surrogate 

aircraft were inserted to replace 737 type aircraft, with a stage length of 800 nm or less, an initial range 

requirement estimate. A production rate of 400/yr was assumed to begin in 2030 and continue for 10 

years, for approximately 4,000 vehicles inserted into the NAS fleet mix. 

Figure 4-15. Impact of N+3 Surrogate on Baseline JPDO 1.8 Times Traffic NAS Total Delay, 

Compared to Baseline Fleet Mix 

At the higher levels of daily flight counts, there is a clear advantage to enabling the metroplex 

operations, and it must be adapted to some extent to hope to reach the predicted demand levels. These 

operations, for the 5,000 foot field length capability, can offload a significant amount of traffic from the 

OEP airports, as shown in Figure 4-16(b). Focus is made in this study on the metroplex operations at the 

OEP airports since these are the major economic centers through which the vast majority of air traffic 

flows, and traffic delays at these airports are highly correlated with delays elsewhere in the system 

[Ref. 29]. 

Figure 4-16. (a) OEP Airport Distribution; (b) Effect of Metroplex Operations on Traffic 

Reductions on OEP Airports 

The current baseline operations model, which assumes typical airport assets, but includes operational 

NextGen improvements, does not offer an acceptable delay level, or conversely, traffic throughput for a 

fixed delay. So various levels of metroplex infrastructure adaptation levels were investigated to determine 

the pareto frontier for NextGen enabled towered or non-towered fields. This was done as a tradeoff, not as 

(a) (b)

From: http://www.faa.gov/about/initiatives/nextgen/

NASA/CR—2010-216798 29
a hard driver of field length, so that most of the additional capacity could be added to mitigate the local 

impact to smaller airports that would likely oppose additional aircraft noise, operations traffic, and 

exhaust emissions. With this motivation, the distribution of NextGen enabled resources was investigated, 

using both towered and non-towered airfields assuming virtual tower capabilities. The geographic 

distribution of the limited set (only fields with 4,000 feet or longer runways) shown in Figure 4-17(b) 

indicates the close correlation of towered fields with population centers in Figure 4-17(a). The two sets of 

non-towered runways, one comprised of 4000 feet to 5000 feet runways, the other 5,000 feet or greater 

runways, show very different qualities in their distribution and relationship to the population. The nontowered 5,000 feet and above airfields have a more even correlation with the population, whereas the 

4000 feet bin airfields show more density east of the continental divide and on the eastern seaboard, 

relative to the western half of the U.S. The N+3 vehicle should leverage airfields across the U.S., and 

driving to a 4,000 feet BFL would be less useful for many of the major west coast markets. For a 

developing point-to-point system, that would be well covered geographically by the 5,000 feet runway 

and greater airfields. 

Figure 4-17. (a) U.S. Population Distribution; (b) Airfield Distribution by Longest Field Length for 

Non-towered Runways, Including >4,000 feet and >5,000 feet Bins, Along With Towered Runway 

4.2.2.3 Field Length Requirement – A cumulative histogram of the airfield resource by longest 

runway length shows this dataset in a quantitative fashion. Assuming that all airfields can be equipped 

with a virtual tower, Figure 4-18 shows the relative addition of airfields as a percentage of the number of 

currently towered U.S. airfields. At 6,000 foot fields or greater, there is already 186% (979 fields) of the 

525 towered fields available. At 5,000 feet or greater, 329% (1728 fields) of the towered fields is 

available. Considering the 35 OEP airports carry most of the traffic, and most greatly affect demand, 

driving the field length requirement to use 5,000 foot runways adds a vast set of infrastructure. 

Considering only the satellite airports within 90 nm of the 35 OEP airports, the number of with 5,000 foot 

runways is 709. Decreasing to a 70 nm radius, the number drops to 491, but compared to the number of 

OEP centers, this is a large capacity boost to accommodate increased demand. Simulations were then 

performed with the 70 nm, 5,000 foot runways in the metroplex areas surrounding the OEP airports, to 

estimate the benefit to meeting traffic demands. The baseline scenario is without additional metroplex 

runways and without the N+3 surrogate, while the offload 3 times scenario indicates the metroplex 

operations simulation with the N+3 surrogate vehicle. The JPDO reference future traffic scenario, a 1.8 

times increase in traffic, was simulated for comparison. 

(a) (b)

From: U.S. Census Bureau

NASA/CR—2010-216798 30
Figure 4-18. Histogram of U.S. Airports (Towered and Non-towered) as a Percentage of 

Towered Airports by Max Field Length, Assuming Implementation of Virtual Towers 

Investigating the three times traffic simulation in the metropolitan Atlanta area, it was observed that 

the addition of the metroplex fields and a vehicle that can take advantage of the infrastructure can reduce 

the traffic levels at the OEP airports significantly. The offloading levels were determined by constraining 

throughput at the OEP fields to 90% of capacity, while allowing up to 40/flights per hour at the metroplex 

offload fields. This throughput was enough to allow the three times traffic levels, but at increased delay. 

Figure 4-19 shows the capacity effect of introducing the N+3 surrogate vehicle in the NAS, looking only 

at the Atlanta metroplex. It was possible to maintain a sustainable traffic throughput at ATL while 

distributing traffic to the available 5,000 foot and greater metroplex resources. Atlanta is a good example 

market (although results were similar elsewhere) since its delay is one of the largest in the U.S. (see 

Figure 4-20 which relies on FAA OPSNET delay data), and it is currently the world’s largest airport by 

passenger volume. This was the main objective of the future traffic simulations: to determine which 

resources to include to enable projected demand. However, the levels of delay in these scenarios need 

further consideration. 

Figure 4-19. Comparison of Baseline and N+3 Metroplex Hourly Operations at 

ATL for Three Times Traffic Levels 

674% 673%

637%

493%

329%

186%

118%

75%

52% 37% 23% 15% 6% 2% 1% 0%

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16

Length of Longest Runway (Thousands of Feet)

NASA/CR—2010-216798 31
Figure 4-20. U.S.-Wide Delay Instances Distribution for January-July 2009 Period (OPSNET) 

The result for reducing severe delay was significant for the 1.8 times traffic levels (JPDO) simulation 

in the metroplex operations model, especially at ATL, LAS, PHX, and EWR. Many other airports saw 

substantial decreases in delay as well. The relative levels of delay across the NAS (shown as vertical bars 

indicating the minutes of delay on average) are compared for the 1.8 times prediction in Figure 4-21. The 

left figure shows the baseline scenario, without the metroplex operations, and without the N+3 surrogate, 

while the right figure shows the results with the N+3 surrogate in place with the metroplex operational 

model. 

Figure 4-21. (a) Baseline 1.8 Times Traffic Delay Distributions; (b) Metroplex Operations Delay 

With >5,000 foot Airfields 

NAS delay for the three times traffic level scenario is shown in Figure 4-22, which reduces average 

delays greatly across the board with the addition of the 5,000 foot runways. Compared to the 1.8 times 

scenario, the three times simulation results show a large number of areas with substantial delay while 

maintaining more fields much nearer to their capacities full-time as seen in Figure 4-20. This is very near 

the threshold of requiring a lower field length to include additional resources, but since the tradeoff 

between enabling a larger number of the smaller metroplex airfields, and imposing new aircraft noise and 

emissions seems unclear, it was decided to use 5,000 feet as the driving requirement for field length. This 

(a) (b)

NASA/CR—2010-216798 32
is a direct cause of the high traffic demand levels predicted in the 2030 timeframe, in the range of 1.8-3 

times 2008 levels. Potentially for a later EIS date, it would be necessary to drive to a 4,000 feet, or 

perhaps a 3,000 feet field length requirement. 

Figure 4-22. (a) Baseline Three Times Traffic Delay Distributions; (b) Metroplex Operations Delay 

With >5,000 foot Airfields 

4.2.3 Cruise Speed Sensitivity 

In a related NAS simulation study, the effect of varying the cruise speed to accommodate a 737 

replacement N+2 aircraft was considered using that previously composed dataset [Ref. 30]. This study 

was leveraged to produce a first-order sensitivity of delay resulting from cruise speed changes for a 737-

like aircraft. The simulation decreased the mean cruise speed slightly by approximately 12 knots in the 

manner indicated in Figure 4-23, which shows the speed distribution of the flight simulation set of 

approximately 10,000 flights in and out of the NYC metroplex. The mean cruise speed for the 737 was 

448 knots, while the N+2 vehicle cruised at approximately 436 knots, mostly by shifting the fastest flights 

into the lower speed bins, resulting in a more pronounced peak in the set distribution about the mean 

flight speed. 

Figure 4-23. Cruise Speed Distribution Shift for an N+2 NAS Simulation 

Compared to Baseline 737 Distribution 

The outcome of the speed reduction was an expected effect on delay, acting to increase mean delay 

over the baseline case by about 9.2 minutes on average. That corresponds to an increase of 7.6 minutes of 

delay per airframe for every 10 knots decrease in cruise speed. The delay increase distribution is shown in 

(a) (b)

0

1000

2000

3000

4000

5000

6000

228 248 268 288 308 328 348 368 388 408 428 448 468 488 508

Flight Counts

Knots

Flight Data Set Cruise Speed Distribution

CESTOL

B737

NASA/CR—2010-216798 33
Figure 4-24. This delay level is on par with the average delay for the JPDO 2040 demand set at the 

1.8 times traffic level. This is clearly an unacceptable increase in delay, approximately doubling the 

projected future delay levels, which would in turn reduce traffic throughput. 

Figure 4-24. Aircraft Delay Distribution for a Cruise Speed Decrease of 12 Knots 

Compared to 737 for an N+3 Vehicle Study 

A second implication is that increased delay then 

requires more airframes to complete the same number of 

flights per day. Referring to Figure 4-12, it is clear that 

operating efficiency is dependent on the number of 

segments completed per airframe per day, and that also 

leads to a larger fleet requirement. For the N+3 analysis, 

the effect of cruise speed on the number of required 

airframes was analyzed for a range of Mach numbers 

from 0.60 to 0.85. For a flight replacement strategy 

consisting of replacing 4,000 airframes with ranges of 

less than 1600 nm, without any shift to the metroplex 

concept at the JPDO 1.8 times levels, the effect of cruise 

speed on the number of airframes required is shown in 

Figure 4-25. These flights have airborne cruise times on 

average of only 25 minutes shorter for the Mach 0.85 flights compared to the 0.60 flights. 

This is another advantage of sizing a vehicle to leverage the metroplex concept – that the sensitivity 

to the cruise speed is lower than compared to a large, long-haul transport, and so the effects of slight 

decreases in cruise speed can be mitigated. However, the effect on the number of airframes required, from 

an economic standpoint, is not small even for these relatively short flights investigated. For a Mach 0.60 

mission, there is about an 8% increase over the 0.85 Mach cruise, which is a significant economic driver 

also to keep cruise speed high. The cruise speed requirement was set at Mach 0.75 or higher in order to 

maintain traffic levels and the current delay averages, and to minimize the economic impacts of cruising 

at potentially slower speeds to improve vehicle-level efficiencies. This lower bound represents a slight 

decrease from the 737 average speed of about Mach 0.77 for the same mission mix seen in the baseline 

JPDO demand set. The vehicle optimization trades will be allowed to approach this lower bound, but not 

venture lower to preserve the ability of the platform to meet the top level objective of delivering projected 

future traffic levels. 

4.2.4 Passenger Loading and Range Capabilities 

To best size the payload (PAX count) and range capabilities of the N+3 type vehicle, a study of the 

current vehicle fleet mix was used to determine the most broadly applicable airframe entry into the future 

-

200 

400 

600 

800 

1,000 

1,200 

1,400 

1,600 

1,800 

-21-18-15-12

-9-6-3

0369

1215182124273033363942454851545760

Flight Counts

Gate Arrival Delay Difference (minutes)

Statistics (minutes)

Figure 4-25. Effect of Varying Cruise Mach 

Number on Number of Airframes Required 

to Complete 8,000 Flights 

NASA/CR—2010-216798 34
landscape for use with the metroplex concept. The current fleet mix was investigated since the 

extrapolation of usage for future flight levels is very similar, and detailed demand studies were not 

performed to investigate additional point-to-point routes enabled by the N+3 type vehicle between 

metroplex satellite airport pairs. The Terminal Area Forecast (TAF) was used to determine and then 

further extrapolate the future fleet mix as indicated in Figure 4-26. Although there is some variation 

between the growth rates of individual vehicle classes, the most prominent airframes persist to the future 

demand scenarios, being the sub-30, 60-90, and 120-150 passenger classes. The philosophy for the 

vehicle insertion into the commercial fleet was to assume a production rate of 400/yr, and perform the 

NAS simulations in 2040 when 4,000 vehicles have been inserted. Based on this level, and to take 

advantage of metroplex operations without an insurmountable requirement for a high-lift system to make 

the 5,000 foot field length requirement (and also to achieve high loading rates), the primary need and best 

tradeoff appeared to be at the 120-150 passenger level. 

Figure 4-26. Flight Count Projections for Various Future Daily Passenger Traffic Growth Levels 

To refine the passenger level further, the distribution of most often flown vehicles currently in use 

was investigated with respect to range capability to determine if there was a clear need to fly many shorter 

routes typical of smaller RJ type vehicles, or whether there was a need to deliver a larger passenger count 

on a longer cruise typical of some B757 or A320 aircraft. Figure 4-27 shows the number of flights for 

each airframe currently in use in the CONUS, as well as the typical range capability usage. The majority 

of the airframes with the most use rarely exceed 1,600 nm more than 25% of the time. The correlation 

with Figure 4-26 is clear, in that it serves the best need of the current and extrapolated demand 

distribution to bias toward a smaller vehicle in the 120-150 seat class, and so a nominal value of 120 

passengers was adopted for the vehicle requirement. 

-

10,000 

20,000 

30,000 

40,000 

50,000 

60,000 

70,000 

Number of Flights

Number of Seats per Airframe

2025

2040

2086

NASA/CR—2010-216798 35
Figure 4-27. Number of Flights per Airframe Type (Line) Superimposed on Range Capability 

Statistical Bars for U.S. Aircraft 

To preserve similar operating characteristics 

throughout the range of most-used aircraft the N+3 

vehicle could potentially replace, the vehicle range 

capability of 1,600 nm was also adopted. Of the current 

flight segment distribution for CONUS flights, this 

clearly positions the airframe to perform the vast 

majority of operations as seen in Figure 4-28. This 

ensures that the broadest need segment will be served 

by the advanced technology platform, while also 

bolstering compatibility with metroplex operations in 

order to meet traffic levels. 

4.2.5 Mission Requirements 

The goal of this requirements study was to 

determine the proper vehicle capabilities to meet the 

broad, primary need of the future scenario. Future 

traffic projections were developed over a range 

determined by JPDO estimates as well as the most 

challenging estimates of the scenario analysis, from 1.8 

times levels to approximately 3 times levels. 

Simulations of delay and traffic performance with NextGen implementation alone, compared to a fleet 

containing an N+3 vehicle that enabled the metroplex, were performed. For both the 1.8 times and 3 times 

levels, the N+3 metroplex concept greatly reduced delay compared to the baseline simulation with 

NextGen only. A study into existing runways revealed that the ability to use non-towered airports can 

dramatically increase the number of available runways for use in the 2030-2035 timeframe, and thus 

makes the metroplex concept viable. Using non-towered airports with 5,000 feet or greater runways can 

add enough alternative fields to keep pace with projected traffic growth, although delay will suffer at the 

higher traffic projections. Furthermore, a bypass alternative study at ATL showed that the integration of 

the N+3 vehicle into the NAS while taking advantage of these airports can reduce delay times 16% for 

low traffic days and up to 44% in cases of high demand. For this reason, and the preceding discussions, it 

was decided to implement a mission requirement of a 5,000 ft field length for the aircraft. A maximum 

wingspan of 150 ft was adopted to accommodate the space constraints associated with metroplex 

Figure 4-28. Cumulative Histogram of Flight 

Counts by Stage Length in the Current 

CONUS NAS 

nm 

NASA/CR—2010-216798 36
runways. To mitigate the impact on operational efficiency, it was decided to allow a small decrease in 

cruise Mach number down to 0.75. This will help ensure the N+3 vehicle does not exacerbate delay any 

further within the traffic ranges predicted. An important assumption was made that these traffic levels will 

not greatly alter the future distribution of average fleet capabilities, in terms of number of operations at 

the various stage lengths and passenger loadings. Thus, the passenger and range capabilities were set to 

meet the vast majority of operations in the projected fleet mix, which is similar to the current picture. A 

summary of the requirements determined in this study is given in Table 4-3. 

Table 4-3. N+3 Mission Requirements 

Mission Requirements

Range (with reserves): 1600 nm

Passengers: 120

Balanced Field Length (Sea Level/Standard Day): 5000 ft

Landing Distance (Sea Level/Standard Day): 5000 ft

Minimum Cruise Mach: 0.75

NASA/CR—2010-216798 37
5 ADVANCED VEHICLE CONCEPTS 

5.1 Engine Architecture Concepts 

Three advanced engine configurations 

projected into the 2030-2035 EIS timeframe were 

developed, the major features of which will also be 

described here. At an interim point in the sizing of 

the aircraft, re-sized engines were generated based 

on updated thrust requirements. These re-sized 

engines also incorporated advanced propulsion 

technologies to be described in the sections below. 

Selection of engine architectures for initial 

development were driven by the outputs of the 

technology QFD study (Figure 5-1). 

5.1.1 Reference Engine 

The CFM56-3B1 turbofan used on the 737-500 was used as a 

scalable reference engine (Figure 5-2). It was necessary to acquire 

engine data for the CFM International CFM56-3B1 turbofan in 

order to quantify the mission fuel consumed by the reference 

vehicle, and hence to develop the N+3 fuel burn target. This dualshaft turbofan uses a 60 inch diameter fan consisting of 38 

titanium blades with a maximum speed of 5,490 RPM. The axial 

low-pressure (LP) and high-pressure (HP) compressors contain 

three and nine stages, respectively, comprising an overall pressure 

ratio of 27.5 at maximum climb. The fully annular combustor 

leads into a single-stage HP turbine followed by a four-stage LP 

turbine. The fan duct achieves a bypass ratio of 6.0, leading to a maximum takeoff thrust of 20,000 lbf, a 

cruise thrust of 4,650 lbf, and a cruise specific fuel consumption of 0.67 pph/lbf. 

5.1.2 Three-Shaft Turbofan 

The 2030 EIS three-shaft advanced technology turbofan uses separate shafts for the fan module, 

intermediate pressure (IP) compressor, and a combination axial/centrifugal HP compressor. Three-shaft 

turbofans are a proven technology and are in broad application in the Rolls-Royce Trent engines 

[Ref. 31]. A mechanical profile diagram of the engine is shown in Figure 5-3. At takeoff, this engine 

achieves a bypass ratio near 18, with a fan pressure ratio of 1.34 and an overall pressure ratio (OPR) of 

50. At SLS conditions, the engine achieves a specific fuel consumption of 0.226 pph/lbf. The high flow 

rate fan module consists of 22 composite fan blades combined with 54 swept stator vanes. In terms of 

installation effects, the engine model assumes a constant core nozzle Cv of 0.995, no customer bleed, fan 

stream suppression generated by the core nozzle which drives it to a base Cd of 0.995, and a constant 

power extraction of 100 hp. An inlet recovery of 99.7% is also assumed. 

Figure 5-1. Engine Downselection Process 

Figure 5-2. CFM56-3B1 Turbofan 

3-Shaft Turbofan Open Rotor Geared Turbofan

Open Rotor 3-Shaft Turbofan

CFM56-3B1

NASA/CR—2010-216798 38
Figure 5-3. Three-Shaft Turbofan Mechanical Profile Diagram 

Technologies embedded inside the engine include active tip clearance control in the HPC and high 

pressure turbine (HPT), high overall pressure ratio (OPR) materials and seals, and active flow control in 

the integrated power system (IPC). Additional advanced high temperature materials are employed, fuelcooled cooling air is used to reduce ceramic matrix composite turbine blade temperatures, and highstrength shaft materials provide additional structural integrity. The engine also uses high temperature 

sumps, electric fuel and oil pumps, and an embedded starter/generator. Lean combustor technology 

enables low NOx production. Finally, shape memory alloys enable a variable geometry (two-position) 

nozzle for reduced takeoff jet velocity and noise. At a TRL of 6, this engine technology is expected to be 

fully mature by the N+3 timeframe. Fuel consumption characteristics of the engine will be discussed in 

Section 5.1.5. 

5.1.3 Geared Turbofan 

An engine architecture receiving widespread attention is the geared turbofan [Ref. 32]. Pratt & 

Whitney is anticipating the certification of its PW1000G geared turbofan in 2011, preferring the noise 

reduction capability of this configuration over the fuel consumption benefits of the open rotor [Ref. 33]. 

Similar in design to the three-shaft turbofan, the geared turbofan incorporates an additional design 

consideration in the gearbox which breaks the direct shaft connection between the LPC/LPT and the fan 

module. This allows for both the fan module and the LPT to be operating in their optimum speed range, 

instead of one being constrained by the other. Originally estimated at a TRL of 5, recent flight testing of 

the PW1000G on the 747SP and the A340 has led to an increased TRL of this propulsion system, now 

estimated to be in the range of 6-7 [Ref. 34]. While the engine performance far surpasses that of the 

CFM56-3B1, its fuel consumption is nearly identical to that of the three-shaft turbofan. At SLS 

conditions, the engine achieves a specific fuel consumption of 0.229 pph/lbf. In terms of installation 

effects, a constant core nozzle Cv of 0.995 is assumed, along with no customer bleed, fan stream 

suppression generated by the core nozzle which drives it to a base Cd of 0.995, and a constant power 

extraction of 100 hp. An inlet recovery of 99.7% is modeled. 

Figure 5-4. Geared Turbofan Mechanical Profile Diagram 

5.1.4 Open Rotor 

The open rotor engine concept has been revived in recent years by most major engine producers due 

to increasing fuel costs [Ref. 35]. Previous research conducted by General Electric on the GE36 Unducted 

NASA/CR—2010-216798 39
Fan and Pratt and Whitney/Allison on the 578-DX similarly followed high-fuel price trends in the 1970s 

[Ref. 36]. Though the open rotor exhibits the high fuel efficiency inherent in turboprop engines, its thrustto-weight ratio is significantly less than that of the three-shaft turbofan (Figure 5-5). The open rotor 

achieves a takeoff SLS thrust-to-weight ratio of 2.32 with respect to the dry engine system, whereas the 

three-shaft turbofan reaches up to 6.15. During cruise, the difference is not quite as significant, with the 

open rotor and three-shaft turbofan achieving thrust-to-weight ratios of 0.47 and 0.87, respectively. These 

weight differences are due to the massive nature of the propeller system. When comparing to the three 

shaft turbofan, the weights of each common module are quite similar. It is the addition of the gearbox and 

propeller pitch change mechanism that penalizes weight so significantly. First, the gearbox outputs 

enormous torques due to the slow rotational speed of the propellers. This torque level requires the size of 

the gearbox to be significant as a percentage of engine weight. In addition to the gearbox, the pitch 

change mechanism is another reason for the open rotor weight penalty. Each individual propeller blade 

requires independent and redundant pitch control mechanisms. These mechanisms must overcome the 

centrifugal loading from 7 foot propeller blades while precisely controlling attack angles to tenths of a 

degree. Significant acoustic penalties also accompany this configuration due to its external counterrotating propellers. While some aeroacoustic prediction capabilities have been developed for counterrotating propellers by Whitfield et al [Ref. 37,38] and Hanson [Ref. 39], validation of these tools with 

experimental data has been limited. High-fidelity numerical models have not yet been developed in the 

public domain. 

Figure 5-5. Open Rotor Mechanical Profile Diagram 

The 2030 EIS open rotor engine uses an IP compressor and a combination axial/centrifugal HP 

compressor. Hot-side flow is used in conjunction with a gearbox to drive the counter-rotating propeller 

system. At SLS takeoff the engine achieves an OPR of 35 and a specific fuel consumption of 0.157 

pph/lbf. The model for the engine core assumes a constant nozzle Cv of 0.995, no customer bleed, and a 

constant power extraction of 100 hp. Ideal inlet recovery is also assumed. 

Technologies embedded inside the engine include active tip clearance control in the HPC and HPT, 

high OPR materials and seals, and active flow control in the IPC. Advanced high-temperature materials 

are employed, fuel-cooled cooling air is used to reduce turbine blade temperatures, and high-strength 

shaft materials provide additional structural integrity. The engine also uses high-temperature sumps, 

electric fuel and oil pumps, and an embedded starter/generator. Lean combustor technology enables low 

NOx production. Currently estimated at a TRL of 5, this technology is expected to be fully mature by the 

N+3 timeframe. 

5.1.5 Engine Comparison and Downselection 

Figure 5-6 shows a comparison between the specific fuel consumption of the geared turbofan and the 

three-shaft turbofan when sized for the same takeoff thrust. It can be seen that the performance of the 

engines is nearly identical. At idle, the three-shaft turbofan has a slightly improved fuel consumption 

compared to that of the geared turbofan. Similarly, Figure 5-7 compares the specific fuel consumption of 

the advanced engine architectures compared to the reference CFM56 engine. It can be seen that the 

NASA/CR—2010-216798 40
advanced technology engines exhibit significantly reduced fuel burn compared to the baseline engine 

(44% and 60% reduction for the three-shaft turbofan and open rotor, respectively, during this portion of 

the mission). While this comparison is a huge selling point for the open rotor architecture, it does not 

articulate either the difference in weight or noise between the two engines. Again, the geared turbofan 

showed the same characteristics as the three-shaft turbofan and, as a result, is not displayed on this graph. 

The geared turbofan was set aside for the remainder of the study, not because of any deficiency in the 

technology, but rather to limit the number of possible combinations of configuration and propulsion 

system. It is expected that any results stemming from the study of the three-shaft engine will be applicable 

to any aircraft using a high-bypass ratio turbofan architecture, including a geared turbofan. 

Figure 5-6. Geared Turbofan and Three-Shaft Turbofan SLS 

SFC Hook Comparison at Comparable Technology Level 

Figure 5-7. Engine Architecture SLS SFC Hook Comparison, Advanced Technology Level 

5.2 Airframe Concepts 

The airframe concept downselection process is shown in Figure 5-8. Beginning with a wide initial 

design space, various analyses were made to reduce the number of possible configurations for detailed 

study. The aircraft comprising the initial design space will be described below, followed by a description 

of the processes which narrowed the field of candidate configurations. 

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0 5000 10000 15000 20000 25000

Specific Fuel Consumption, pph/lbf

Net Thrust, lbf

Geared Turbofan

Three-Shaft Turbofan

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0.50

0 5000 10000 15000 20000 25000

Specific Fuel Consumption, pph/lbf

Net Thrust, lbf

Three-Shaft Turbofan

Open Rotor

CFM56-3B1 Turbofan

Idle Max T/O

NASA/CR—2010-216798 41
Figure 5-8 .Airframe Configuration Downselection Process 

5.2.1 Tube-and-Wing 

The tube-and-wing vehicle configuration is the classic design 

configuration for modern transport aircraft. This configuration offers 

good packaging for passengers, cargo, and propulsion systems while 

maintaining relatively low drag characteristics (Figure 5-9). The engine 

placement on the undersides of each wing permits low-cost, 

straightforward integration and easy maintenance access to the propulsion 

system. The large main wing is balanced in pitch by the smaller rear 

horizontal tail, which can be placed at any height along the vertical tail. 

When the engines are placed in an aft-mounted configuration, a T-tail 

empennage is required to minimize the interference of the engine exhaust 

with the control surfaces. A study of this alternate configuration is 

presented in Section 6.1.1. However, T-tails result in heavier vertical tails out of structural necessity 

[Ref. 40]. The inverse proportionality of aspect ratio to induced drag suggests that this is a primary design 

driver of low fuel consumption aircraft. However, the additional wing weight that accompanies larger 

wings must be taking into consideration. The cylindrical geometry of the fuselage required for passenger 

seating forces a fineness ratio in the range (greater than three) associated with low fuselage friction drag 

[Ref. 41]. While not a revolutionary configuration, advanced technologies that may be mature by the 

2030-2035 EIS timeframe may unlock advanced versions of this configuration by allowing for higher 

aspect ratio wings and reduced drag through the implementation of surface features on the fuselage or 

laminar flow wing design. Such configurations will be referred to as advanced tube-and-wings (ATWs). 

Among the disadvantages of this configuration is the drag associated with the wetted area of podded 

engines and aft empennage [Ref. 42]. 

Initial Design Space

Qualitative Down-select

Quantitative Down-select

Use of reliever 

airports

r 

Figure 5-9. Tube and Wing 

Configuration 

NASA/CR—2010-216798 42
5.2.2 Tube-and-Wing With Embedded Turbofans

The tube-and-wing configuration with an embedded 

propulsion system offers the same advantages as the tube-andwing configuration with potential for noise reduction benefits 

due to the acoustic shielding inherent in embedded engines 

(Figure 5-10). Furthermore, there is a potential reduction in 

aircraft wetted area due to the removal of the engine pods, 

depending on the extent of engine integration. An example of 

this type of configuration in the commercial industry is the de 

Havilland Comet 4 [Ref. 43]. Powered by four Rolls-Royce 

Avon Mk 524 turbojets, each wing contained two engines 

embedded near the root. However, this configuration 

experienced undesirable levels of fuel consumption. The 

implementation of the buried inlet introduces losses in pressure recovery, resulting in a reduction in the 

propulsive efficiency capabilities of the engine. Additionally, placing the engine near the wing root can 

lead to the ingestion of boundary layer flow off of the fuselage, further hindering the pressure recovery 

potential of the inlet. Since the propulsive efficiency of the engine is inversely proportional to specific 

fuel consumption, both inlet pressure recovery losses and boundary layer prohibit the engine from 

achieving its full potential. The flow over the wing in the vicinity of the engine inlets can become 

disturbed and lead to a reduction in lift capacity, counteracting the wetted area reduction achieved by 

removing the engine pods [Ref. 42]. Additionally, modern commercial transport speeds and their 

respective wing sweeps would result in high spillage drag with wing-embedded propulsion systems 

[Ref. 44]. While additional technologies could be included to reduce inlet pressure distortion and alleviate 

flow disturbances over the wing, these technologies would add weight to the system, on top of fuel weight 

that must compensate for reduced engine efficiency. In addition, active systems may require substantial 

amounts of engine bleed. Perhaps the most compelling argument against this type of aircraft configuration 

is the incompatibility of this design with the use of a very high bypass turbofan. In order to embed a 

large-diameter engine inside the wing, the volume must be available at the wing root. However, because 

of the large diameters associated with fans, this becomes impractical. Furthermore, the open rotor engine 

architecture is not feasible with this type of aircraft design. Due to the penalties significantly outweighing 

the benefits, tube-and-wing embedded propulsion configurations were not considered for further analysis. 

5.2.3 Tube-and-Wing With Canard 

The addition of a lifting canard onto a conventional tube-and-wing 

configuration with or without aft tail geometries provides an alternative 

method to achieve stability and control. Its implementation in the 

absence of an aft horizontal tail results in a more unconventional 

geometry to control (Figure 5-11). Work by McGeer and Kroo has 

shown that this type of configuration will generally result in a higher 

total drag than the conventional layout [Ref. 45]. Canard designs studied 

by Keith and Selberg showed that the wing loading on the canards 

required to trim the aircraft was excessive, leading to extreme levels of 

induced drag [Ref. 46]. Vertical placement of the canard has been 

shown to be a main design driver behind this type of configuration. 

When the canard is combined with an aft horizontal tail, minimum control surface deflections can be used 

which minimizes trim drag [Ref. 47]. However, additional weight penalties will be realized, in addition to 

the increased skin friction and induced drag arising from the presence of another lifting surface [Ref. 48]. 

Figure 5-10. Wing-Root Embedded 

Turbofans With Canard 

Figure 5-11. Tube and Wing 

Configuration With Canard 

NASA/CR—2010-216798 43
5.2.4 Hybrid Wing-Body

The hybrid wing-body is a revolutionary alternative to the tubeand-wing configuration considered in this concept design study 

(Figure 5-12). Reports have shown that the hybrid wing-body 

(HWB) configuration can have significantly decreased root bending 

stresses compared to traditional tube-and-wing configurations 

[Ref. 49]. 

This reduction in bending loads, combined with the vehicle 

operating at a significantly lower wing loading, allows for increased 

structural efficiency. Additionally, decreased wing loading would 

contribute to a decrease in takeoff and landing field length. HWB 

vehicles also permit the elimination of tail sections. Eliminating 

these surfaces would decrease vehicle drag through a reduction in 

skin friction and form drag, increasing lift-to-drag and reducing fuel burn [Ref. 50]. The centerbody of the 

HWB allows for engine noise reduction by acting as a noise shield. In a study by Papamoschou et al 

[Ref. 51], a cumulative effective perceived noise level (EPNL) reduction of 7.4 EPNdB was possible 

through the use of centerbody noise shielding with an aft shield length of five jet diameters. The study 

suggests that increased exhaust mixing would increase the effectiveness of noise shielding. In addition, 

boarding operations become faster and more efficient due to the seating layout of the vehicle, decreasing 

turnaround time during ground operations. 

Though this configuration offers structural benefits, it presents a structural challenge in creating a 

pressure vessel for the cabin without adding significant empty weight. The configuration deviates from 

the traditional cylinder vessel to a pentagon-shaped design, which is more challenging to design with a 

low weight and ability to support adequate loads during cabin pressurization [Ref. 49]. Past studies have 

concluded that hybrid wing-bodies might be best used for large cargo or passenger missions, as opposed 

to shorter mission lengths and passenger counts due to the internal volume of the vehicle increasing faster 

than the external surface area. For large vehicles and passenger counts, this could lead to decreased drag; 

however, scaling the vehicle for a smaller passenger count results in the internal volume decreasing faster 

than the external surface area which would lead to an increase in drag [Ref. 50]. 

The minimum centerbody height is determined by the size of the passengers, regardless of the 

mission. Centerbody volume is subsequently set by passenger count. This is shown graphically in Figure 

5-13. Wing spar and engine location then drive the length of the fuselage, driving the fuselage thicknessto-chord ratio. As a result, the minimum allowable size for the hybrid wing-body is dictated by passenger 

count, regardless of the mission. Based on the passenger count defined in the mission requirements, the 

minimum fuselage length was 70 feet, further constraining the root chord of the outboard wings. 

Figure 5-13. HWB With Embedded Propulsors Configuration 

6 Foot Height 

Constraint

Aft Spar Location

Home Plate Area Used 

to Calculate Number of 

Passengers

Plan

View

Closure Angle

Constraint

Figure 5-12. HWB With 

Podded Propulsors 

NASA/CR—2010-216798 44
5.2.5 Hybrid Wing-Body With Embedded Turbofans

The HWB configuration with an embedded propulsion 

system offers the same advantages as the HWB configuration 

with the additional benefit of further engine noise reduction 

(Figure 5-14). This noise reduction would be a result of engine 

noise shielding due to the burying of the engines into the body. 

Integrating the propulsion systems would reduce wetted area due 

to the removal of the nacelles, assuming volume is available in 

the centerbody, and would allow augmentation of the vehicle 

aerodynamics through the use of the propulsion system [Ref. 50]. 

Furthermore, reductions in ram drag can be realized due to the 

ingestion of boundary layer flow into the inlets. 

The introduction of buried inlets introduces pressure recovery penalties that can translate into losses 

in propulsive efficiency. Specifically, embedding the propulsion system increases inlet distortion, 

worsening the efficiency of the system compared to a podded engine configuration. This decrease in 

propulsive efficiency would result in increased fuel burn [Ref. 52]. Embedded engines would also occupy 

internal volume in the fuselage which would otherwise be available for passengers and cargo. With a 

fixed passenger count, adding internal volume to support an embedded propulsion system would lead to 

an increase in wetted area, empty weight, and drag. In addition, embedding the propulsion system would 

add complexity to engine maintenance and decrease vehicle structural volume [Ref. 50]. Due to the 

penalties relative to the N+3 mission goals significantly outweighing the benefits, HWB embedded 

propulsion configurations were not considered for further analysis. 

5.2.6 Channel Wing 

An adaptation of the tube-and-wing concept, the channel wing 

consists of a fuselage and aft tail with modified wing geometry 

(Figure 5-15). Instead of a flat wing planform, a circular channel is 

built into the wing in which the engines are mounted. The vehicle 

can use the same variety of aft tail geometries common on tube-andwing configurations, and employ a canard if deemed necessary. The 

implementation of the engine channels allows for the increase in 

velocity over the upper surface of the channel to augment the 

circulation and lift of the wing [Ref. 53]. While stall angles have 

been shown to be extended out to 45 degrees using a channel wing, the high lift associated with these 

extreme angles of attack have been inaccessible due to tail scrape angle limitations. Hence, channel wing 

geometries are typically accompanied by circulation control to achieve lift coefficients in the range of 8.5 

to 9.0 [Ref. 54]. Channel diameter will be constrained by propeller size, and may lead to a reduced thrust 

potential of the open rotor because half of the propeller disk is effectively ducted. An experimental study 

by Blick and Homer showed that this penalty could be as much as 10% [Ref. 55]. As the basic geometry 

remains in the shape of a tube-and-wing, the same benefits can be sought in this configuration by 

maximizing aspect ratio to minimize induced drag and seeking laminar flow over the flat portions of the 

wing. However, it must be taken into consideration that smaller portions of the wing are available for 

laminar flow purposes. The channel wing configuration has increased risk associated with it over more 

conventional designs due to the limited amount of channel wing research. Compared to more 

conventional platforms, the channel in the wing also presents a structural challenge that must be 

addressed. Furthermore, a flat wing of the same projected platform will mathematically produce less 

profile drag for a lower structural weight. 

Figure 5-14. HWB With Embedded 

Propulsors Configuration 

Figure 5-15. Channel Wing 

Configuration With Canard

NASA/CR—2010-216798 45
5.2.7 Joined Wing

The joined wing configuration consists of a typical cylindrical 

fuselage with a swept front wing and a vertical tail. Instead of a 

horizontal stabilizer, a second forward-swept wing, whose root is 

typically placed at some spanwise position along the length of the 

vertical tail, joins the trailing edge of the front wing, as shown in Figure 

5-16. It has been shown that a joined wing can weigh 22%-35% less 

than an aerodynamically equivalent cantilevered wing and tail 

[Ref. 56,57]. This has been attributed to the rear wing behaving as a 

strut for the front wing. Wolkovitch showed in his work that using 

trailing edge control surfaces should be sufficient for all stability and 

control requirements [Ref. 58]. 

In addition, the joined wing configuration distributes the wing area over a larger portion of the 

vehicle, decreasing wave drag and improving transonic aerodynamic properties [Ref. 56]. However, as a 

result of using a second intersecting wing, excess wetted area would exist and interference drag would 

increase from surface intersections. In addition, joined wing vehicles are more difficult to trim at 

maximum lift coefficients for equivalent tube-and-wing configurations [Ref. 59]. As seen in Figure 5-16, 

like the tube-and-wing, the joined wing configuration would offer no engine shielding from the main 

wing, diminishing its ability to contribute to the N+3 acoustic goal. 

5.2.8 Low Aspect Ratio Spanloader 

Low-aspect ratio spanloader (LARS) configurations offer 

improved boarding operations due to the seating configurations and 

entrances offered by a spacious passenger cabin layout (Figure 

5-17). With the ability to load and unload passengers more quickly, 

reduced time associated with ground operations can be realized by 

this vehicle design. This would improve the time utilization of the 

vehicle when introduced into NAS operations. This configuration 

also offers a viable platform for distributed propulsion and boundary 

layer-ingesting inlets. However, similar to the issue encountered 

with hybrid wing-body aircraft, using a non-cylindrical fuselage 

introduces difficulties in pressure vessel design. For a rectangular 

fuselage with the same order of magnitude in perimeter as a round 

fuselage, the associated bending stresses can be an order of magnitude higher [Ref. 60]. Low-aspect ratio 

experimental aircraft have shown that these configurations can be built extremely light, reducing their 

required lift and, in turn, the associated induced drag [Ref. 61]. However, the low aspect ratio nature of 

the vehicle will counteract these induced drag benefits. Furthermore, it is expected that these studies 

would not translate well into a full-sized commercial transport, as the size of the vehicle will be driven by 

the size of the passenger cabin. 

Though LARS was selected as a candidate platform for studies of distributed propulsion, it was 

determined that many of the enabling technologies behind this form of propulsion system would not be 

ready for production given the N+3 timeframe. However, because the configuration held promise in terms 

of ground operations, it was carried through the qualitative down-select process to be examined in more 

depth. 

Figure 5-16. Joined Wing 

Configuration 

Figure 5-17. LARS 

Configuration 

NASA/CR—2010-216798 46
5.2.9 Configuration Quantitative Assessment and Downselect 

In order to downselect the advanced 

body configurations (Sections 5.2.3, 5.2.6-

5.2.3), the vehicles were quantitatively 

assessed with respect to five parameters 

determined to be representative of the 

mission goals: aerodynamic drag, mission 

fuel, empty weight, vehicle noise, and field 

length. For each of these metrics, the 

airframe configurations were assigned 

scores based off of their estimated 

performance with respect to one another. 

These scores were subsequently converted 

into a figure of merit that shows the relative 

benefits and disadvantages of the 

configurations, shown in Figure 5-18. For example, because of the large wetted area and drag associated 

with the LARS configuration, it received a low score in the aerodynamic drag category. Due to the 

potential weight savings associated with the joined wing configuration, a high score was assigned to its 

empty weight category. The hybrid wing-body and the channel wing were determined to have the best 

potential to shield engine noise. With this scoring system, and with these metrics, this initial quantitative 

assessment between airframe configurations yielded the hybrid wing-body, channel wing, and tube-andwing vehicles as the three most promising airframes to be carried forward. LARS configuration was 

determined to be the least promising of the airframe configurations, followed by the joined wing and the 

canard configurations. 

Less weight was placed on extreme short 

takeoff and landing capability as a result of the 

analysis of field length requirements. Unusually 

high lift coefficients were not needed to achieve 

metroplex operations. Traditional high lift 

systems proved to be sufficient. Relaxing the 

field length constraint in the vehicle qualitative 

assessment yielded the tube-and-wing and 

hybrid wing-body as the two preferred 

configurations out of the six vehicles, as seen in 

Figure 5-19. In addition to these quantitative 

results, the added complexity and risk associated 

with the channel wing led to the dismissal of the 

configuration from further consideration. The 

result of the quantitative assessment showed that 

the conventional nature of the tube-and-wing configuration may be the most suitable platform upon which 

to build and mature future technologies. Furthermore, for the N+3 EIS timeframe, it may be the best 

balance of risk and solution for the mission objectives. A summary of the configuration downselection 

process is illustrated in Figure 5-8. 

5.3 New Technology 

The assessment of candidate technologies for aircraft entering service in the N+3 timeframe required 

careful consideration of the relative benefits and penalties associated with each technology, as well as 

their potential for maturation in a timeframe consistent with their incorporation into a new aircraft. In 

order to perform this analysis, an extensive matrix of candidate technologies was developed, as described 

in Section 5.3.1. This list of technologies was then qualitatively ranked according to a QFD process, 

Figure 5-18. Configuration Quantitative Assessment 

Figure 5-19. Configuration Quantitative 

Assessment Discounting Field Length 

NASA/CR—2010-216798 47
described in Section 5.3.2. Technologies consistent with the N+3 timeframe were selected from among 

the top candidates. The remainder of this section details the reasoning behind the selection of certain 

technologies for continued study. It will further describe the models that were used to depict the benefits 

and penalties of those technologies that were chosen for further examination. A diagram of the process is 

shown in Figure 5-20. 

Figure 5-20. Technology Assessment Process Flow 

5.3.1 Development of Technology Matrix 

In the generation of the original technology matrix, input was gathered from subject matter experts 

(SMEs) on the Northrop Grumman team. Each group submitted technology concepts within its area of 

expertise, e.g., an extensive list of candidate propulsion and noise technologies was developed in 

collaboration with Rolls-Royce Liberty Works. Appendix A lists the candidate technologies broken into 

four major subheadings: airframe, aerodynamics, noise, and propulsion. 

5.3.2 Quality Function Deployment Process 

In order to rank the candidate technologies in preparation for the final downselection, a QFD method 

was used. QFD made it possible to translate qualitative assessments of the relative benefits and penalties 

of individual technologies into a common quantitative foundation. SMEs familiar with the technologies 

were called upon to rate each technology with respect to the N+3 objectives. These scores were used to 

develop technology effectiveness ratings (TERs), or scenario-weighted rankings of each technology. Five 

potential scores were used to define the impact of each technology with respect to the N+3 goals. For 

each goal, the SME could assign a value of 1, 3, 5, 7, or 9, where a value of 3 represented a technology 

having no impact on a specific N+3 goal, a value of 1 corresponded to a detrimental impact, and 5, 7, and 

9 were relative levels of positive benefits. When weighted for each scenario, the summation of these 

scores comprised the TER for the technology. Input was also gathered from SMEs concerning TRLs for 

all candidate technologies. A snapshot of the master QFD worksheet at an interim point in the information 

gathering process is shown in Figure 5-21. A full listing of the QFD worksheet is given in Appendix D. 

NASA/CR—2010-216798 48
Figure 5-21. Format of Master QFD Worksheet and Database 

In order to properly accommodate for low-TRL technology risk in the technology effectiveness 

ratings, a TRL Risk Factor was applied. The curve, shown in Figure 5-22, emphasizes how technologies 

currently rated below TRL 3 are penalized through this risk factor, while those greater than TRL 3 are 

given advantage due to their higher likelihood to mature in the N+3 timeframe. 

Figure 5-22. Applied TRL Risk Index Factor 

The final weighting parameter considered in the TER framework is the relative importance of each of 

the N+3 metrics in the overall scenario defined in Section 4.1.2. Since fuel burn and emissions were the 

metrics of most concern to the projected future scenario, technologies that most impacted these metrics 

positively were given additional weighting in the TER system. Although not considered in the overall 

TER for each technology, interactions between technologies were assessed in a separate interaction 

quotient, to illuminate mutually beneficial and exclusive technology packages. For the overall scenario, 

Figure 5-23 shows a compiled TER worksheet which was used in the technology down-selection process, 

taking into consideration the risk index factor. The top ten technologies for each scenario are shown in 

Figure 5-24. 

SELECT Goals:

SER 

Total

Scale Factor (Priority)

Normalized Scale Factor

Units:

Goal

Threshold

Fz (Threshold Value On Gaussian Curve):

Max Score This is the maximum score 9 25.00 9 25.00 9 25.00 9 25.00 100.0 9 10 5 9 2

Baseline Baseline 3 7.50 3 7.50 3 7.50 3 7.50 30.0 1 100 0 1 0

ID Designation: Description: Physical Principles:

FB 

Input

FB 

SER

FL 

Input

FL 

SER

NOx 

Input

NOx 

SER

LDN 

Input

LDN 

SER

Total 

SER

TRL 

Input

TRL 

Risk

Interact

ion 

IQ 

Input

IQ 

Score Comments / Assumptions

92 Prop Concept #4.24 TO/C-Assist Propulsion Unit

Detachable takeoff/climb autonomous 

propulsion units propel vehicle to cruise 

altitude, then return to airport increasing cruise 

efficiency. 5 14.64 9 25.00 5 14.64 5 14.64 68.9 1 100.0 2,3,4

Assume lower flight velocity for low-power 

"engine" matching. Decreased cruise weight.

39 Noise Concept #3.11

Ultra-High (10+) Bypass Ratio 

Turbofans Trading Velocity for Mass Flow Rate 7 21.87 5 14.64 3 7.50 7 21.87 65.9 6 65.7 3,4

High bypass allow for large mass flow @ sea 

level, and improved SFC with small weight and 

penalty despite nacelle drag.

56 Noise Concept #3.28

Wing Morphing (Shape Memory 

Alloys, Piezoelectrics) High-Lift Device Integration into the Wing 3 7.50 7 21.87 5 14.64 7 21.87 65.9 2 95.0 1,2,3

Assumes high lift benefit without cruise penalties 

typically associated w/ fixed high lift devices

20 Aero Concept #2.06

Distributed Exhaust Nozzle Flap 

(DEN Flap)

Distributed Exhaust Nozzles Integrated Into 

Flap Provide Flow Control/Supercirculation. 1 2.94 9 25.00 5 14.64 7 21.87 64.5 2,3,4

Hi lift systems typically add weight and increase 

cruise drag, but greatly improve takeoff 

performance.

21 Aero Concept #2.07

Steady Circulation Control (Offtake 

Engine Flow Used for Circulation 

Control via Blown Flap, IBF, 

Coanda Jet Flap, etc.)

Low-Cm Exhaust Provides Flow Control and 

Circulation Control Wing (CCW) 1 2.94 9 25.00 5 14.64 7 21.87 64.5 2,3,4

Hi lift systems typically add weight and increase 

cruise drag, but greatly improve takeoff 

performance.

22 Aero Concept #2.08

Unsteady Circulation Control 

(Offtake Engine Flow Used for 

Circulation Control via Blown Flap, 

IBF, Coanda Jet Flap, etc.)

Pulsed Effectors Provide Similar Benefits as 

Steady CCW 1 2.94 9 25.00 5 14.64 7 21.87 64.5 2,3,4

Hi lift systems typically add weight and increase 

cruise drag, but greatly improve takeoff 

performance.

35 Noise Concept #3.07 Geared Turbofan Allows Very High-Bypass Cycles 5 14.64 1 2.94 9 25.00 7 21.87 64.5 5 75.0 1,3,4

Weight Penalty, moderatley better SFC, 

moderate to high reduction in NOx.

Multiple options available. Could be combined 

Interaction Quotient

1.00

TRL

0.25

3

25

Fuel Burn

1.00

Qualitative

9

Take-off & 

Landing Field 

Length

NOx 

Emissions Noise

1.00 1.00 1.00

25 25 25

Qualitative Qualitative Qualitative

3

9 9 9

3 3

0.3 0.3 0.3 0.3

NASA/CR—2010-216798 49
Figure 5-23. Technology Effectiveness Ratings for Overall Scenario 

Figure 5-24. Top Ten Technologies for Each Scenario Based on QFD Analysis 

Technology Effectiveness Rating

Magnitude by TER for Overall Scenario, sorted by TER w/ RI

0

10

20

30

40

50

60

70

80

90

100

35 70 36 102 27 56 21 1 5 9 22 100 72 90 24 95 30 58 88 41 15 16 34 28 74 80 87 83 12 29 18 25 75 96 60 38 44 31 63 85 62 53 32 52 37 48 55 3 11 14 54

Technology ID Number

Fuel Burn Field Length

LTO NOx Noise LDN

Geared Turbofan

Ultra-High (10+) Bypass Ratio Turbofans

Constant Volume Combustion / Wave Rotors

#1

#2

#3

#4 Open Rotor with Counter-Rotating Fan 

#5 Variable Geometry Nozzles

#6 Turbo-Electric Distributed Fans

#7 Non-Turbine Distributed Fans

# DESCRIPTION TER 

Rating*

1 Geared Turbine 85.3

2 Variable Geometry Nozzles 79.0

3 Non-Turbine Distributed Fans 77.8

4 Steady Circulation Control 69.6

5 Wing Morphing 67.4

6 Distributed Exhaust Nozzle (DEN) Flap 66.9

7 Unsteady Circulation Control 66.9

8 Ultra-High Bypass Ratio Turbofan 64.8

9 Hybrid All-Electric Propulsion 63.2

10 Noise Shielding 62.6

# DESCRIPTION TER 

Rating*

1 Ultra-High Bypass Ratio Turbofan 67.8

2 Steady Circulation Control 65.1

3 Wing Morphing 63.3

4 Geared Turbine 63.1

5 Distributed Exhaust Nozzle (DEN) Flap 62.5

6 Unsteady Circulation Control 62.5

7 Open Rotor 61.9

8 Modeling for Engine Inlet Optimization 58.6

9 Inlet Flow Control 58.6

10 Variable Geometry Nozzles 58.1

rbine Distributed Fans

Circulation Control

orphing

uted Exhaust Nozzle (DEN) Flap

bine

Exhaust Nozzle (DEN) Flap

rculation Control

# DESCRIPTION TER 

Rating*

1 Geared Turbine 74.8

2 Ultra-High Bypass Ratio Turbofan 67.9

3 CVC / Wave Rotor 66.2

4 Open Rotor 65.5

5 Variable Geometry Nozzles 63.1

6 Turbo-Electric Distributed Fans 61.9

7 Non-Turbine Distributed Fans 60.7

8 Modeling for Engine Inlet Optimization 58.6

9 Inlet Flow Control 58.6

10 Turboshaft-Powered Direct Drive 

Distributed Fans

58.3

Wave Rotor

Rotor

e Geometry Nozzles

Electric Distributed Fans

# DESCRIPTION TER 

Rating*

1 CVC / Wave Rotor 91.9

2 Open Rotor 88.9

3 Geared Turbine 76.0

4 Turbo-Electric Distributed Fans 74.3

5 Ultra-High Bypass Ratio Turbofan 70.9

6 Swept-Wing Laminar Flow 68.9

7 All Electric Engines (no AGB) 67.4

8 Microvortex Generators / Riblets 67.4

9 Turboshaft-Powered Direct Drive 

Distributed Fans

64.7

10 Deployable Vortex Generators 60.9

Geared Turbine 76.0

Turbo-Electric Distributed Fans 74.3

Ultra-High Bypass Ratio Turbofan 70.9

Swept-Wing Laminar Flow 68.9

Rank

Rank

Rank

Rank

NASA/CR—2010-216798 50
5.3.3 Down-Selection of Candidate Technologies 

While the QFD process succeeded in sorting the candidate technologies best suited for the overall 

future scenario, it was necessary to reduce the number of technologies to a manageable number for more 

detailed investigation. Those technologies receiving the highest TER rating were considered first, and 

divided into the disciplines of aerodynamics, propulsion, structures, and noise. Decisions were made to 

limit the number of technologies among the disciplines to more equally distribute the candidate 

technologies. In this section, technologies selected for further consideration will be discussed, and 

modeling within the framework of the design tools will be described in detail. Technologies were 

modeled in such a way that they could be packaged to accurately reflect the additive effects of combining 

individual technologies. As technologies are packaged together, the net system benefit increases. 

However, the benefits do not necessarily combine in a linear fashion. For example, if two separate 

technologies both reduce vehicle empty weight by 50%, adding these technologies would not result in a 

100% decrease in empty weight. Coupling these technologies would yield a net reduction less than 

100%, but still greater than the 50% reduction seen by each individual technology. 

Technologies that were considered but set aside are listed, along with a discussion of the rationale, in 

Appendix B. Technology maturation plans for the preferred technology package for the final vehicle 

design, along with current TRL level estimates and required development milestone paths to reach TRL 6 

by 2025 are provided in Chapter 7. For those technologies that are already beyond TRL 6, no additional 

development path was developed, and it is assumed that sufficient investment will occur to make these 

available for the N+3 EIS. 

5.3.3.1 Natural Swept-Wing Laminar Flow – Regions of laminar flow greatly reduce profile 

drag through skin friction drag reduction compared to those of turbulent flow [Ref. 62]. Implementation 

of laminar flow can lead to a significant reduction in vehicle drag and a subsequent reduction in fuel burn. 

For a typical transonic commercial aircraft, natural laminar flow over 25% of the upper and lower 

surfaces would result in an approximately 25% reduction in wing profile drag [Ref. 63]. Natural sweptwing laminar flow is enabled through the use of airfoils conducive to laminar flow, elimination of surface 

breaks, and a more favorable surface geometry. A laminar flow tool was developed to calculate and 

visualize the potential regions of laminar flow on a given surface. In-house experimental research has 

quantified the achievable chord-wise extent of laminar flow regions in the form of maximum transition 

Reynolds number as a function of wing sweep, and these empirical results have been incorporated into 

this model. Based on this proprietary research and experimental testing, the maximum transition Reynolds 

number was found for the given sweep and projected to the 2030 timeframe. The geometry of the surface 

is input, along with vehicle operating conditions such as altitude and Mach number. The tool then plots 

the regions of laminar flow on the given surface, along with calculating the percentage of the surface in 

the laminar flow region, as in Figure 5-25. The tool takes into account turbulent wakes caused from 

section interference, such as the wing-fuselage interface, engine and nacelle, pylon, wing tips, and control 

surfaces. No regions of laminar flow can exist aft of section breaks, such as flaps. The drag divergence 

Mach number of the vehicle was corrected, as the presence of shocks precludes the existence of laminar 

flow. In addition, the maximum lift coefficient is penalized due to slat removal and airfoil profile change 

to accommodate laminar flow. 

NASA/CR—2010-216798 51
Figure 5-25. Laminar Flow Regions Visualization for High Aspect Ratio Wing 

The percentage of laminar flow per surface, the corrected drag divergence Mach number, and the 

maximum lift coefficient penalties were modeled in the aerodynamics module in FLOPS. The existence 

of laminar flow in MIDAS was reflected by the removal of slat noise during acoustic calculations. 

5.3.3.2 Landing Gear Assembly Component Integration – Landing gear assembly component 

integration entails the incorporation of smaller landing gear assembly components into more inclusive 

parts. This may often include placing smaller components inside 

larger components, reducing landing gear drag by removing 

sources of small-scale vortex shedding. Consequently, reduced 

turbulent and vortex shedding leads to a reduction in source (and 

far-field) noise during approach. Landing gear must be designed 

to be structurally sound, and the integration of smaller 

components must be performed in a manner that does not 

sacrifice the sturdiness of the gear. A large downside to this 

process is the difficulty in maintenance that would result from 

reducing access to smaller landing gear components. 

According to Guo et al [Ref. 64], a clean landing gear 

configuration is attenuated, conservatively, by 3 dB of broadband 

noise compared to a fully dressed landing gear configuration for 

wind tunnel Mach numbers between 0.18-0.24. Furthermore, 

referring to Quayle et al [Ref. 65], a faired landing gear sees a 

drag reduction of approximately 30%. According to in-house 

SMEs, an integrated assembly would not be quite as 

aerodynamic as a faired gear, so the estimated drag reduction 

was estimated to be 25%. 

Eqn. 6

Landing gear component integration was modeled in FLOPS as a drag reduction specific to the 

landing gear in the aerodynamics module, at no cost to weight for either the main and nose landing gear. 

In MIDAS, landing gear component integration was modeled as a landing gear part count reduction. 

5.3.3.3 Landing Gear Fairings – Landing gear assembly fairings are designed to streamline the 

landing gear assembly during landing and takeoff of the aircraft (Figure 5-26). Vortex shedding due to the 

bluff body nature of both the nose and main landing gear is a major contributor to high amplitude 

airframe noise. Fairings are a passive method by which to present a more aerodynamic surface to the 

oncoming flow during low altitude operations, at which distances to observers are minimal. Fairings can 

be applied to any landing gear as long as geometric constraints allow for them. However, they will add 

extra weight and possibly volume to the system. Furthermore, they must be designed to deploy 

Nacelle Turbulent Wedge

Engine Centerline

Nacelle Turbulent Wedge

Laminar Flow Region

Control Surfaces

Wing Tip

Turbulent Wedge

Fuselage Turbulent Wedge

DGear DGear C 0.25C

Figure 5-26. Landing Gear Fairing 

NASA/CR—2010-216798 52
simultaneously with the landing gear assembly, adding complexity to the overall system. Reduced drag 

during landing associated with the addition of fairings is typically undesirable to pilots, as landing gear 

drag aids in slowing down the aircraft. However, in light of the potential acoustic benefits, this has not 

been taken into consideration. Failure of a landing gear assembly fairing could cause damage in many 

ways. The possibility of it detaching from the aircraft during takeoff and landing would pose a threat to 

neighborhoods surrounding airports. For an aircraft, the landing gear assembly that loses its fairing would 

generate high noise and drag. Furthermore, the loss of the fairing would result in asymmetric drag loading 

on the landing gear. At most directivity angles, Dobrzynski et al experimentally found a broadband noise 

reduction of approximately 2.7 dB [Ref. 66]. A landing gear drag reduction study predicts approximately 

30% reduction in landing gear drag due to the addition of fairings [Ref. 65]. Weight penalties are 

estimates based on landing gear size. 

Eqn. 7

Eqn. 8

Landing gear fairings were modeled in FLOPS as a landing gear drag reduction and a weight 

increase. The landing gear drag coefficient in FLOPS was corrected to reflect the aerodynamic effects of 

fairings in a manner similar to the modeling of landing gear component integration. To be conservative, 

landing gear fairings were modeled in MIDAS as a broadband reduction of 2 dB from the source sound 

pressure level (SPL) of the landing gear at all directivity angles. Landing gear fairing technology is wellestablished, at a current TRL of 8, but is not prevalent in modern commercial aircraft. However, the 

extreme nature of the N+3 acoustic goal is expected to revitalize interest in this technology, and drive it 

towards designs more widely accepted in the industry. 

5.3.3.4 Distributed Exhaust Nozzle (DEN) and Flap – The distributed exhaust nozzle concept 

involves the transition from a typically round exhaust nozzle to a nozzle with a large array of smaller 

holes through which the engine exhaust may flow. These smaller holes are generally rectangular or 

triangular. Because the exhaust area becomes restricted (even if it matches that of a baseline round 

nozzle), there is a thrust penalty associated with distributed exhaust nozzles. However, the noise reduction 

potential is significant. Exhausting through small holes instead of a large area creates smaller jets with a 

higher wave-number peak in the turbulence spectra, shifting the far-field noise to higher frequencies 

where it can be attenuated by the atmosphere to a greater extent [Ref. 67]. This technology could possibly 

be retrofitted to an existing exhaust nozzle, but in practical applications it would replace an existing 

nozzle. It would need to be designed to the same temperature constraints as existing nozzles, though there 

may be additional pressure requirements because of the flow restrictions imposed by the porous exhaust 

surface. Creating a nozzle from an array of small holes leads to the potential of clogging or the melting 

closed off some of the exhaust ports. If this were to happen, the overall thrust would diminish, but there is 

also a high degree of redundancy in the system with so many holes in the nozzle. If the structure between 

a series of holes were to rupture, the flow through the nozzle would tend towards this larger opening 

(possibly causing an even larger rupture), and there would be a source noise increase associated with this 

larger flow. In Gaeta et al [Ref. 67], a 10% reduction in thrust is reported due to DEN, based on 

computational results. A 4-8 dB reduction in far field noise is achieved at subsonic jet velocities, with an 

increase in high frequency noise due to the shifting of the peak frequency. Extrapolating this concept to a 

flap system, engine exhaust flow is diverted through ducting to the trailing edges of the wings. This flow 

is then distributed evenly across the wing span, and blown over the trailing edge flap system and resulting 

in increased circulation and lift. 

The DEN flap was modeled in FLOPS as an increase in lift coefficient during takeoff and landing, as 

well as a thrust penalty across the entire flight envelope. The increase in lift coefficient is a function of the 

deflection angle of the flap and the percentage of the span that experiences blowing as seen in Eqn. 10. 

The thrust penalty is also modeled as a function of flap deflection angle, as seen in Eqn. 9. 

DGear DGear C 0.3 C

WGear 0 2. WGear

NASA/CR—2010-216798 53
Eqn. 9

Eqn. 10

Eqn. 11

Eqn. 12

where is the flap deflection, bblown in the percentage of the span over which the nozzle exhausts, and 

b is the total span. 

The engine weight parameter in the FLOPS weights module was corrected to accurately reflect the 

increase in engine weight as a result of installing DEN flaps. There is a drag penalty associated with the 

technology that was applied in the aerodynamics module across the entire mission. DEN was modeled in 

MIDAS as a reduction in sound pressure level for all frequencies associated with jet noise and flap 

sources. Currently estimated at TRL 2, this technology should mature quickly due to the similarity of its 

subsystems to those of circulation control technology. 

5.3.3.5 Boundary Layer Ingesting Inlet – Boundary layer ingesting (BLI) inlets are a propulsive 

concept that could lead to an increase in propulsive efficiency. The engines are partially embedded while 

still having a portion of the propulsion system protruding into the flow near the vehicle surface. This 

allows for the ingestion of a segment of the boundary layer that has formed along the vehicle [Ref. 68]. 

Engine exhaust normally produces an increased region of velocity aft of the vehicle, while lowmomentum boundary layers produce a decreased region of velocity aft of the vehicle, leading to a 

nonuniform velocity profile in the wake. Nonuniform velocity profiles aft of the vehicle lead to a decrease 

in propulsive efficiency [Ref. 107]. Ingestion of this low-momentum flow allows for the production of a 

more uniform velocity profile aft of the vehicle and a reduction in ram drag, which improves propulsive 

efficiency, as illustrated in Figure 5-27. However, due to the embedded inlet and inlet placement in the 

boundary layer, inlet distortion is increased compared to podded engine configurations. Inlet distortion 

decreases pressure recovery, which consequently decreases propulsive efficiency. To measure the 

effectiveness of this technology, reduction in drag and increased inlet distortion were modeled in terms of 

aerodynamic and propulsion, with no consideration given to packaging and integration. Results from 

Rodriguez were scaled for a two-engine configuration and used for analysis [Ref. 69]. BLI inlets on the 

hybrid wing-body were approximated to reduce vehicle zero-lift drag 2.2% and pressure recovery 2.9%. 

Eqn. 14

Eqn. 13

where pr represents pressure recovery. A percentage decrease in inlet pressure recovery decreases 

engine thrust by approximately the same percentage. BLI inlets were modeled in FLOPS as a drag 

reduction in the aerodynamics module. This module in FLOPS contains lift-induced drag and zero-lift 

drag coefficient parameters. The zero-lift drag coefficient parameter was corrected to reflect the reduction 

in drag for two engines. Due to the increase in inlet distortion from BLI inlets, a thrust penalty was 

applied to the entire engine table prior to being input into FLOPS. For commercial applications, BLI 

inlets are estimated to be at TRL 2. However, the technology is expected to mature quickly due to the 

groundwork laid by its application in military vehicles. 

Tfactor cos 0. 15

b

b C blown

L 15

Wengine 18Wengine 0.

CD 0, .001

0, 022 0, CD .0 CD

pr 029pr 0. pr

NASA/CR—2010-216798 54
Figure 5-27. Aft Velocity Profile for Ideal BLI and Podded Propulsion Configurations 

5.3.3.6 Two Degree-of-Freedom Liner – Two degree-of-freedom (DOF) liners are built upon the 

current state-of-the-art single DOF acoustic liners for reducing tonal fan noise. Double DOF liners 

contain two layers of absorbing material and additional perforated sheets separating them from the other 

layers. Similar to single DOF liners, two DOF liners can have their absorbing sections divided up into 

smaller cells designed to target specific tonal frequencies. Two DOF liners allow attenuation to target two 

different tonal frequencies. A tradeoff exists between the absorption capacity of the liner and the drag that 

is generated by the porous face sheet. Additionally, the weight and volume penalties associated with 

acoustic liners increase as the degrees of freedom increase. Based on analysis, it was found that a two 

DOF inlet liner provided superior noise reduction and overall system performance compared to a single 

DOF liner and multi-DOF liners. Two DOF acoustic liners are currently estimated to be TRL 8, and this 

technology was determined to be the best solution for inlet fan noise suppression. Analysis and results of 

the liner design study are described in Section 6.1.2. 

5.3.3.7 Deployable Vortex Generators – Deployable vortex generators (VG) are used to increase 

the lift of an aerodynamic surface while minimizing the drag penalty. They consist of small ramps of 

various geometries that are typically scaled to the size of the boundary layer. By generating vortices in the 

flow, the boundary layer stays energized and separation is subsequently delayed [Ref. 70]. In this manner, 

higher lift capabilities exist and reduced noise is achieved. Vortex generators can be applied to slats, 

flaps, or wing surfaces, and may be either active or passive. In a passive system, they must be affixed to 

the aerodynamic surface, penetrate through the aircraft skin, or be built into the airframe itself. An active 

system will involve a deployment capability of the vortex generators, leading to weight penalties, but no 

additional drag penalties at cruise conditions. Failure of this technology could involve either the 

detachment of a vortex generator from the wing surface, or the failure of the deployment mechanism. If a 

vortex generator were to become detached from the wing surface or the deployment failed, the designed 

acoustic and aerodynamic benefits would not be achieved, and airframe noise would be increased. If the 

vortex generator actuator were to fail in the deployed state, drag penalties would be incurred at all flight 

conditions. It must be noted that the specific application of this technology in this study targets noise 

associated with slats. Estimated sound pressure level reductions were taken from Kuo and Sarigul-Klijn 

describing a retractable VG system and its influence on noise [Ref. 71]. Although VGs and microtabs are 

technically different, they are equivalent from an acoustic and aerodynamic standpoint. In terms of weight 

estimates, piezo-compliant linear actuators by Flexsys, Inc. exist on the order of 20 grams per actuator 

(preliminary description can be found in Osborn et al [Ref. 72] with subsequent specifications of the 

constructed system on the Flexsys website [Ref. 73]), and have been used in the construction of 

deployable vortex generator models. An efficient VG array will have approximately nine VG’s per 2 feet 

Podded

NASA/CR—2010-216798 55
of wing span, so assuming a conservative 40 grams per actuator to accommodate wiring and installation, a 

formal relationship between wingspan and weight penalty can be developed: 

Eqn. 15

where b is the wingspan over which the vortex generators are applied, is their linear density along 

the wing, and WVG is the weight of a single vortex generator with allowances for installation effects. 

Deployable vortex generators were modeled in the FLOPS weights module by increasing the wing weight 

parameter to reflect an increase in weight due to their installation. Vortex generators were modeled in 

MIDAS as a narrowband reduction in sound pressure levels in the range of 1-3 kHz to target slat noise 

sources. Furthermore, higher frequency noise penalties were assessed to account for the shifting of noise 

to these frequencies. Estimated at TRL 5, this technology is fully expected to mature in time for 2030-

2035 EIS commercial transport aircraft. 

5.3.3.8 M5 Ultrahigh-Performance Fiber – M5 ultrahigh-performance fiber is a high-strength, 

low-density polymer fiber, prepared from monomers by the addition of an extremely hydrophilic and 

acidic compound to align the monomers and fuse them through chemical bonds along the length of the 

fibers [Ref. 74]. The fibers undergo hydrogen bonding, a process absent in previous composites. This 

bonding between the fibers and graphanic planes allows for a significant increase in strength. In addition, 

M5 fiber exhibits impressive thermal and flame resistances. M5 could be substituted for many, if not all, 

composite applications on most current aircraft, such as the Boeing 787, serving as a replacement for the 

commonly-used carbon fiber-reinforced plastic. M5 can be integrated into the fuselage structure, main 

wing, both vertical and horizontal tail sections, and nacelle. M5 fiber would allow for larger integrated 

structures, thus reducing part count and further reducing structural weight. Limitations of application 

include leading edges of the wing, tails, and nacelles, which historically use metallic materials, as these 

surfaces often encounter debris. Weight reduction estimates, based on the report by Cunniff et al, suggest 

that current M5 fiber technology exhibits a strength of 3.96 GPa, a failure strain of 1.4%, and a modulus 

of 271 GPa [Ref. 74]. Conservative estimates of the properties that can be achieved with M5 fiber in the 

near future are a strength of 8.5 GPa, a failure strain of 2.5%, and a modulus of 300 GPa. According to 

Cunniff, the ultimate goals are 9.5 GPa, 2.5%, and 450 GPa, respectively. The density of M5 fiber is 

1.7g/cm3

. For reference purposes, the strength, modulus of elasticity, and density of carbon fiberreinforced plastic are 3.0 GPa, 220 GPa, and 1.71-1.72 g/cm3

, respectively. The significantly improved 

structural properties stemming from the use of M5 fiber, in addition to its relatively low density, would 

result in potentially large weight reductions. The percentage of vehicle structure that could be replaced by 

M5 fiber is equivalent to that of a state-of-the-art aircraft, whose airframes are approximately 50% 

composite by weight [Ref. 84]. 

Eqn. 16

where Wi corresponds to component structural weight, with i representing the components to which 

this structural benefit can be applied: wing, tail, fuselage, and nacelle structures. If this technology is used 

in combination with advanced metallics, described in Section 5.3.3.13, the relative contribution of this 

technology to structural weight reduction must be decreased. M5 ultrahigh-performance fiber was not 

modeled explicitly in MIDAS, as it is strictly an airframe weight reduction technology. Currently 

estimated at TRL 3, this technology is predicted to mature sufficiently for use in the development of 

commercial transports entering service in the 2030-2035 timeframe. 

Wwing bWVG

Wi 22.0 Wi

NASA/CR—2010-216798 56
5.3.3.9 Integrated Aeroservoelastic Structures – Aeroservoelasticity uses directional stiffness in 

the aircraft structural design to control aeroelastic deformation in such a fashion as to affect the 

aerodynamic, control, and structural performance in a beneficial way [Ref. 75]. Performance 

characteristics are a function of many parameters, including weight, aerodynamics, flight mechanics, 

dynamics, and structures. All of these categories should influence wing design from an early conceptual 

stage in the design process. Multidisciplinary design optimization (MDO) methods can be used to design 

a wing for specific loads, flight conditions and performance, improving design efficiency. MDO and 

aeroservoelastic structures use structural deformation of lifting surfaces to achieve performance 

objectives not associated with structural design and find the most favorable design point that satisfies 

structural, weight, aerodynamic, performance, and control constraints. MDO allows the wing of an 

aircraft to experience significant weight reduction as a result of the structure being designed to specific 

conditions and restrictions through gust/flutter/maneuver load control. Since aerodynamics is factored 

into the conceptual design process, aerodynamic loads are better distributed, resulting in a decrease in 

drag. Loads on control surfaces can 

be decreased, resulting in reduced 

power requirements, and hence 

actuation system weight 

[Ref. 76,77]. Based on internal 

research, and data from Wakayama 

and Kroo, flight control system, 

wing, and tail weight reductions 

were estimated as a function of their 

respective aspect ratios as shown in 

Figure 5-28 [Ref. 77]. The aspect 

ratios of the horizontal tail, vertical 

tail, and wing were used to compute 

the respective weight reduction from 

aeroservoelastic structures. Higher 

aspect ratio surfaces exhibited greater deflection than lower aspect ratio surfaces, allowing for increased 

aeroservoelastic tailoring, and thus greater weight reduction. 

 Eqn. 17

Eqn. 18

where fi corresponds to the weight reduction multiplication factor as a function of the respective 

aspect ratio and Wi corresponds to component structural weight, with i representing the components to 

which this structural and subsystem benefit can be applied: wing, horizontal tail, and vertical. 

Aeroservoelastic structures were modeled in FLOPS as a weight reduction in these particular parts of the 

aircraft, but were not modeled explicitly in MIDAS, since the technology is primarily an airframe weight 

reduction technology. However, aeroservoelastic design could potentially reduce control surface size, 

thereby affecting control surface noise. Noise from high-lift systems, landing gear, and engines are much 

more dominant than the noise produced by control surfaces, and so no effort was made to model this 

effect. 

5.3.3.10 Affordable Airframe Large Integrated Structures – Affordable airframe large integrated 

structures allow for a unitized structure methodology yielding specifically tailored structures for given 

loads and load paths. This leads to an increase in structural efficiency, reductions in manufacturing 

processes, time, and cost, as well as a drastic reduction in part count. This technology eliminates 

structural discontinuities and fastened assemblies, while increasing structural efficiency [Ref. 78]. Large 

integrated structures use new design capabilities and manufacturing options, coupled with advancements 

in metallic alloys, composites, and composite joint technology, to achieve a new level of design flexibility 

in unitized structures [Ref. 79]. Developments in materials reduce cost and weight, while new design 

methods further reduce structural weight. Affordable large integrated structures would replace current 

ii AR Wi W f )(

WSurfaceControls WSurfaceControls 0.09

Figure 5-28. Surface Weight Reduction Through 

Aeroservoelastic Structures as a Function of Aspect Ratio

0.0%

2.0%

4.0%

6.0%

8.0%

10.0%

12.0%

14.0%

16.0%

0 2 4 6 8 10 12 14

% Surface Weight Reduciton

Aspect Ratio

NASA/CR—2010-216798 57
primary and secondary vehicle structures in order to create more unitized structures, and offer a weight 

benefit: 

Eqn. 19

where Wi corresponds to component structural weight, with i representing the components to which 

this structural and subsystem benefit can be applied: wing, tail, fuselage, and nacelle structures, in 

addition to the fuel and plumbing subsystem. Affordable large integrated structures were modeled in the 

FLOPS weights module as an airframe weight reduction. Large integrated structures were not modeled 

explicitly in MIDAS. 

5.3.3.11 3-D Woven/Stitched Composites – 3-D 

woven/stitched composites facilitate larger, integrated 

structures by replacing traditionally metallic joints on 

highly-loaded primary structures. Composite pi-joints, as 

depicted in the diagram of Figure 5-29, allow for the 

creation of highly-loaded joints without large and heavy 

metallic fasteners. This is possible due to the joint 

providing a large surface area for bonding a substructure 

to the structure skin. This joint technology allows for 

complex preforms to be connected and fused with one 

other, reducing part count and allowing for complex 

structures to be designed and integrated into the vehicle. 

Three-D stitching exploits the orthotropic properties of 

carbon fiber and limits out-of-plane failure modes through stitching. This allows for failure arrest design 

in the structure. All of the materials in stitched and woven assemblies are dry. This eliminates out-time 

and autoclave requirements, which are normally found in prepreg systems and limit component size. 

These composite joints would replace all of the metallic joints in structural applications, including, but 

not limited to, skin/stiffened fuselage panels and window belts, fuselage circumferential frames, and 

fuselage keel beam frames. On wing structures, this technology can be utilized in joints on the upper and 

lower cover panels, stiffeners and spar caps, intercostal clips, and main landing gear attachments. 

According to Velicki and Thrash, this technology has been integrated into the C-17, including fairings, 

gear doors, and integral fuselage [Ref. 80]. Based on research and testing performed by Sheahen and 

Bersuch [Ref. 81] and internal research, weight reduction from replacement of metal joints with 3-D 

woven and stitched composites was estimated to be 2.62%. This reduction was applied to the wing, tail, 

fuselage, and nacelle structural weights: 

Eqn. 20

where Wi corresponds to component structural weight. Currently estimated to be TRL 3, a roadmap 

has been established for this technology to be sufficiently mature to drive the design of aircraft entering 

service in the 2030-2035 timeframe. 

Wi 05.0 Wi

Wi 0262.0 Wi

Figure 5-29. 3-D Woven/Stitched 

Composite Structural Concept [Ref. 80] 

NASA/CR—2010-216798 58
5.3.3.12 Inverted Flow Nozzle – In order to invert the bypass and core engine streams, a flow 

crossover duct is embedded between the engine nozzle and turbine. By properly shaping the velocity 

profile exiting the engine, a significant degree of noise attenuation can be achieved. Specifically, placing 

the high-speed, high-temperature primary flow on the outside and forcing the low-speed, cooler fan flow 

to the inside, generates two mixing interfaces for the core stream [Ref. 82]. Since more mixing is 

available, noise generation is reduced. Furthermore, noise generated 

at the internal mixing interface is less likely to propagate. This 

technology would need to be incorporated into the original design of 

an engine. The crossover duct would add significant weight to the 

engine system, at a small performance cost. Failure of the 

technology would involve either the failure of the structure of the 

crossover duct, or melting due to the high temperature jet engine 

environment. In either case, there would be a risk of further damage 

to the engine, and the creation of a blockage in the aft portion of the 

engine. 

Experimental wind tunnel measurements of a flow inverter 

nozzle have been conducted at the NASA Ames 40 by 80 foot 

tunnel, shown in Figure 5-30 [Ref. 83]. The inverter tested was 

approximately 36 inches long, embedded inside a 123.5 inches 

JT8D engine. The inverter was 29.1% of the total engine length. 

Although it added weight, it did not add length, as it was embedded 

inside the engine. 

A good first-order approximation is to add 29.1% to the weight of the engine. This is to account for 

the increased volume occupied inside the engine, since the inverter spans the entire engine cross section. 

This work also showed how the overall sound pressure level (OASPL) in the vicinity of the engine was 

affected by the flow inverter nozzle, and gave estimates of the associated thrust penalty. 

Eqn. 21

Eqn. 22

The inverted flow nozzle was modeled in the FLOPS weight module as an increase in engine weight 

as a result of installing a crossover duct into the propulsion system. The weight module contains a 

miscellaneous propulsion weight parameter that is a percentage of the engine weight. This parameter was 

set to a value to accurately simulate the increase in engine weight from this technology. In addition, a 

thrust penalty was applied to the engine deck prior to being input into FLOPS. The technology was 

modeled in MIDAS as a sound pressure level reduction over different frequencies for jet noise source, 

shown in Figure 5-31. 

Wengine 291Wengine 0.

T 0. 985TT

Figure 5-30. NASA ARC 

Inverted Flow Model With 

20-Lobe Mixer [Ref. 83] 

NASA/CR—2010-216798 59
Figure 5-31. Inverted Flow Nozzle Noise Reduction 

5.3.3.13 Advanced Metallic Structural and Subsystem Alloys – New alloys are in high demand 

for structural applications that can withstand elevated temperatures and improve strength-to-weight ratio. 

Advanced metallic alloys could exhibit improved fatigue and crack growth properties. These alloys would 

be used for primary and secondary structural components, as well as subsystem components. The 

increased high-temperature performance of aluminum and titanium alloys also represents significant 

weight saving potential for engines and engine mounts. It was assumed that 45% of the weight of the N+3 

vehicle is metallic weight, similar to the Boeing-787 [Ref. 84]. Internal research has shown that advanced 

metallic alloys offer up to double-digit weight reductions in terms of the percentage of structural weight, 

depending on their application: 

Eqn. 23

Eqn. 24

where Wi corresponds to component structural weight, with i representing the components to which 

this structural benefit can be applied: wing, tail, fuselage, and nacelle structures. If this technology is used 

in combination with M5 ultrahigh-performance fiber, described in Section 5.3.3.8, its maximum relative 

contribution to structural weight reduction must be decreased in Equations 14 and 15. Advanced metallics 

were modeled in the FLOPS weights module as an airframe weight reduction. Advanced metallics were 

not modeled in MIDAS, as it is strictly an airframe weight reduction technology. 

5.3.3.14 Carbon Nanotube Electrical Cables – Carbon nanotubes (CNTs) are manufactured, in 

either sheet or thread form, by injecting fuel and reaction gas into a floating catalyst suspended in a 

furnace. The CNT sheets and threads, which are woven together to form a braid, are extremely 

lightweight, strong, and exhibit electrical conductivity comparable to that of copper wire. A CNT sheet 

can be integrated into a coaxial electrical cable as the replacement for the traditionally copper mesh that 

lies between the outside and inside insulation. CNT woven fiber can be integrated into the coaxial 

electrical cable as the replacement for the traditionally copper wire. Internal research and testing has 

shown significant weight reduction from replacing the copper mesh alone. Replacement of the wire would 

lead to a further increase in weight reduction. 

Eqn. 25

CNT electrical cables were modeled in FLOPS weights module as a subsystem weight reduction. The 

electrical subsystem parameter was reduced in the weights module in FLOPS to reflect the weight 

reduction from CNT electrical cables throughout the vehicle. CNT electrical cables were not modeled in 

MIDAS, as it is a subsystem weight reduction technology. 

0

1

2

3

4

5

6

7

0 20 40 60 80 100 120 140 160 180

OASPL Reduction, dB

Azimuth Angle, deg

i 1.0 WW i .

WLandingGear 1.0 WLandingGear 0.

WElectrical 4.0 WElectrical 0.

NASA/CR—2010-216798 60
5.3.3.15 Modeling for Inlet Optimization – MDO tools and analysis allow for optimum engine 

configuration and propulsive efficiency, while reducing drag in certain body configurations. While 

podded engines on tube-and-wing configurations achieve nearly perfect inlet pressure recovery, hybrid 

wing-body configurations offer potential for further optimization, as their engines are mounted on the 

lifting surface fuselage. According to Rodriguez, MDO analysis for engine configurations on hybrid 

wing-bodies can result in decreased zero-lift drag. However, due to engine configuration changes from 

MDO analysis, inlet distortion may increase. Inlet distortion decreases pressure recovery, and thus 

propulsive efficiency. To reflect the effect of this technology, system-level drag reduction and increased 

inlet distortion were modeled. Results from Rodriguez were scaled for a two-engine configuration and 

used for analysis [Ref. 69]. 

Eqn. 26

Eqn. 27

where pr represents pressure recovery. A percentage decrease in inlet pressure recovery decreases 

engine thrust by approximately the same percentage. The aerodynamics module in FLOPS contains 

induced drag and zero-lift drag coefficient parameters. The zero-lift drag coefficient parameter was 

corrected to reflect the drag reduction for each of the two engines. Due to the predicted increase in inlet 

distortion, a thrust penalty was applied to the entire model. Inlet optimization was not taken into account 

for wing-mounted podded engines due to their sufficiently high level of pressure recovery and inlet 

efficiency. Robust inlet optimization modeling is becoming established as a method by which to 

maximize pressure recovery, and thus propulsive efficiency. 

5.3.3.16 Steady Circulation Control – Circulation control technology extracts engine bleed for 

jet-blowing over a round trailing edge surface to increase lift. A blown sheet of air remains attached to the 

round circulation control (CC) surface, acting as a boundary layer control at low blowing flow rates. At 

higher blowing rates, the blown jet stays attached at the round CC trailing edge, moving the stagnation 

point of the airfoil and its corresponding streamline to the lower surface of the airfoil [Ref. 85]. 

Circulation control technology increases the lift of an airfoil without the need for flaps and other 

mechanical high-lift systems. This high-lift capability enables shorter takeoff and landing distances, as 

well as steeper climb angles. Drag and downstream wakes may also be affected by the implementation of 

steady circulation control as opposed to mechanical high-lift devices. Elimination of mechanical high-lift 

systems reduces the complexity and weight of the wing structure, however this is compensated by the 

ducting required to disperse the engine bleed flow the wing trailing edge. 

Eqn. 28

Steady circulation control was modeled in the FLOPS aerodynamic module as an increase in takeoff 

and landing maximum lift coefficients. The improvement in lift coefficient is given in Equation 28, where 

bblown is the span length over which blowing occurs, and b is the total span of the vehicle. To simulate the 

weight of the engine ducting that would need to be installed, the engine weight parameter in the FLOPS 

weights module was increased appropriately. The engine data table was penalized for thrust to account for 

engine bleed. A drag increase was not included, assuming that an acceptable trailing edge geometry 

(whether active morphing or an innovative integral design) would be implemented. 

5.3.3.17 Engine Shielding – This concept advocates either embedding propulsive components of 

the aircraft into the airframe itself, or locating them above the wings, fuselage, and tails to impede the 

propagation of acoustic waves. This is a well-proven method for reducing aircraft noise, especially at 

higher frequencies, that has found its way into the design of many aircraft [Ref. 86]. Fan shielding alone 

can be accomplished by placing the engine on the top part of the fuselage near the aft section of the 

aircraft, while extensive shielding is achieved by embedding the entire engine into the airframe. 

Utilization of shielding methods must be considered extremely early in the design phase of an aircraft. 

0, 0503 0, CD .0 CD

pr 003pr pr 0.

b

b CC blown C Cll ,max

NASA/CR—2010-216798 61
Based on work from Goodmanson and Gratzer, engine shielding from mounting the nacelles over the 

wing can affect noise significantly, reducing EPNL by approximately 5 dB [Ref. 87]. 

The aerodynamic effects of locating the engine in a position to promote shielding were modeled in 

FLOPS through engine placement. FLOPS allows for the placement of the engines on either the wings or 

aft fuselage. Configuration changes, including tail geometry, were appropriately designed and modeled in 

FLOPS per engine mounting configuration. MIDAS contains routines that enable estimates of noise 

shielding based on the geometry of the aircraft and engine location. 

5.3.3.18 Engine Technologies 

1. Ceramic Matrix Composite Turbine Blades 

Due to their potential for high temperature tolerance, ceramic matrix composite (CMC) turbine blades 

have become increasingly attractive. Their implementation could ease requirements for turbine blade 

cooling, and allow combustors to burn at higher temperatures. CMC turbine blades are also estimated to 

weigh less than their metallic alloy counterparts, contributing to an overall engine weight reduction. 

According to LaChapelle et al [Ref. 88], the use of CMC could increase the temperature range of these 

components between 200-400°R. However, driving the combustor to higher temperatures must be 

balanced against higher emissions levels. At a TRL of 3, it has been projected that this technology will 

have sufficient time to mature before the N+3 timeframe. CMC turbine blades have been assumed for all 

advanced engines under examination in this study. 

2. Lean Staged Combustor 

Staged combustors incorporate multiple distinct combustion zones serviced by independent fuel 

injection systems. These fuel injectors are separated axially in space, and can be optimized at different 

positions along the length of the combustor. Depending on the engine power setting, different 

combinations and power levels of combustor stages can be used. In addition, the burner can be optimized 

to run lean for multiple flight conditions, giving better control over NOx production at all flight 

conditions [Ref. 89]. This is especially significant at the low power settings during which a significant 

portion of LTO NOx is formed. This feature also delivers the capability to reduce velocities in the nearinjector region, reducing acoustic instabilities and improving lean blowout and ignition characteristics. 

While the improvements in LTO NOx can be dramatic for this technology, this benefit must be 

weighed against the added complexity of the fuel system and the associated weight and volume penalties 

that accompany it. 

3. Intercooled Compressor Stages 

This technology involves using a heat exchanger to reduce the temperature of the compressed flow 

between compressor stages. Since it requires less work to compress the lower temperature gas that exits 

the heat exchanger, thermal efficiency gains can be realized. A significant benefit can be achieved with 

reductions in temperature on the order of 40°R. Theoretically, intercooling could be implemented 

between every compressor stage to achieve maximum benefit; however, spatial and weight considerations 

must be made for the heat exchangers. While this technology is currently estimated to be at TRL 3 for 

aircraft applications, it has in fact been implemented on ground-based engines. For example, the LMS100 

ground-based gas turbine engine generates over 100 MW at a thermal efficiency of 46%, in part due to 

the presence of an intercooler between its LP and HP compressors [Ref. 90]. This technology has been 

implemented in the engines under investigation in this study, with intercooler placed between the axial 

and centrifugal stages on the HP spool. Flow losses were under 5%, with upwards of 50% exchanger 

effectiveness by employing endothermic fuel as the heat sink. It was found that implementing the 

technology led to a reduced overall engine weight (including intercools) due to the downstream flowpath 

and component sizing reductions. 

4. Swept Fan Outlet Guide Vanes 

Introducing sweep into the design of the fan outlet guide vanes has the potential to decrease pressure 

losses, while simultaneously reducing fan tip noise. This is achieved through delaying the impact of the 

NASA/CR—2010-216798 62
turbulent wake of the fan rotor onto the leading edges of the outer guide vanes. According to work by 

Schulten, reductions in sound pressure level on the order of 8 dB are possible through vane sweep 

[Ref. 91]. This technology was integrated into the common design of all advanced turbofan engines. 

5. Fan Blade Sweep Design 

Fan blades are swept in the engine to reduce noise and improve engine efficiency. Fan blades operate 

in a transonic environment with the root normally subsonic and the tip supersonic, due to rotational 

velocity. Above relative Mach numbers of 0.8, shock waves begin to form along the tips of the fan blades, 

increasing both noise and shock losses, reducing fan efficiency. Sweeping the fan blades results in a 

reduction of the velocity component normal to the shock surface, weakening shocks and the losses 

associated with their interactions. According to Thomson et al [Ref. 92], forward-swept fan blades may 

provide greater fan efficiency than backward-swept blades by creating a more favorable boundary layer. 

However, forward-swept fan blades are more unstable, as the flow would amplify any vibrations that 

might occur. Conversely, backward-swept blades would dampen vibrations. In addition to poor stability, 

forward-swept fan blades are subject to higher stresses than backward-swept fan blades. In a report by 

Neubert et al, sweeping the fan blades reduced tip Mach number from 1.588 to 0.8, yielding a shock-free 

fan blade and a 1.5% increase in adiabatic fan efficiency [Ref. 93]. Noise levels produced by the fan are 

estimated to be proportional to fan tip speed to the fifth power, which indicates that swept fan blades 

could reduce fan noise significantly. This design feature is assumed in the performance of the advanced 

turbofan engine models. 

6. Lightweight Fan Structure and Cowl 

Through optimization of the front structure of the engine, front-end integration was improved through 

the use of a lightweight fan structures and cowl. This configuration features a greatly shortened fan casing 

which allows for reduction in nacelle length, and thus weight and wetted area, improving aerodynamic 

performance. Highly integrated components lead to engine performance improvements and significant 

weight reduction, including mounting the accessory gear box closer to the core. The lightweight fan 

structure and cowl incorporates a lifted intake and outlet guide vane sweep which further increase engine 

performance and reduce engine noise, drag, and weight. This technology has been assumed in the frontend of the advanced turbofan engines. 

7. Compressor Flow Control 

Extracting diffuser bleed flow, and injecting it into the compressor face in order to mitigate flow 

distortion, may minimize the likelihood of stall or surge. On top of controlling flow distortion, it can also 

be used to dampen flow instabilities that are associated with stall or rotating surge. However, this type of 

flow control requires the use of sensors and controllers to measure, detect, and control instabilities. These 

sensors measure pressure disturbances that occur prior to flow separation, and energize the flow upstream. 

This prevents flow separation from occurring and can increase compressor efficiency. Compressor flow 

control could increase the LPC and HPC polytropic efficiencies by 2% and reduce CO2 emissions by 

3-3.5% compared to current state-of-the-art baseline engines [Ref. 94]. As with any control system, this 

adds weight and complexity to the system, which must be balanced against the potential benefits in 

engine performance. With a current TRL of 4, compressor flow control was modeled into the IPC of the 

candidate engine configurations for this study. 

8. Active Compressor Clearance Control 

The clearance between compressor rotor blade tips and the shroud can vary during flight, due to 

thermal expansion effects. Active compressor clearance control provides higher compressor efficiencies 

by minimizing blade tip losses. This is done by actively maintaining small tip clearances [Ref. 94]. 

Generally taking the form of variable flexible structures, active clearance control can be maintained by 

electromagnetic actuators. However, this type of control requires the use of complex capacitance sensors 

and controllers. This technology is needed to measure, record, and calculate the proper actuator inputs to 

maintain the desired tip clearances. As reported by Weimer, at maximum rotor speed, the blade is only in 

the vicinity of the sensor for approximately 6 milliseconds, requiring exceptionally high frequency 

NASA/CR—2010-216798 63
sensors [Ref. 95]. This experimental work showed about a 1% increase in compressor efficiency imparted 

by active clearance control. Not only does this control system add weight to the engine system, but spatial 

considerations must also be made for such a system. For the engines under investigation in this study, it 

has been implemented in not only the HP compressor, but also in the HP turbine. 

9. Variable Geometry Nozzles 

Variable geometry nozzles use an SMA-actuator to expand and contract nozzle exit area. This 

capability to allows for the control of engine exit conditions, mainly for affecting jet velocity. This allows 

for a trade between engine noise and performance at each power setting and throughout the flight profile. 

This technology is seen on many modern military aircraft to accommodate variable flight conditions and 

performance requirements. Commercial vehicles would use this technology to optimize cruise fuel 

efficiency, as well as reduce noise during takeoff and landing. During takeoff and landing, the engine 

would operate at a designated mass flow rate with a large nozzle exit area and decreased jet velocity, 

reducing noise. This would adversely affect fuel efficiency, but for a very short period in the flight 

profile. When the vehicle reaches cutback altitude where noise is not a concern, the nozzles adjust to 

improve propulsive efficiency for improved fuel burn. The benefit of this technology on propulsive 

efficiency decreases with increasing bypass ratio. Shape memory alloys could allow up to 20% nozzle 

area variability, with a possible 10% increase in nozzle weight. Variable nozzle geometry could also 

reduce CO2 emissions by up to 8.3% [Ref. 94]. 

Jet source noise in MIDAS is calculated using exhaust flow properties at discrete points along the 

aircraft trajectory. 

NASA/CR—2010-216798 64
6 AIR VEHICLE DESIGN STUDY 

Figure 6-1 details the process followed in the synthesis of the air vehicle design which eventually led 

to the selection of the preferred vehicle concept. Airframe configurations were integrated with technology 

packages to form advanced vehicle configurations. These configurations were influenced by trade studies 

on engine mounting location, MDD and sweep, and laminar flow integration. The resulting configurations 

were modeled and analyzed in FLOPS and MIDAS to quantify field length, mission fuel requirements, 

noise characteristics, and emissions performance. A trade study was performed to select an objective 

function for the optimizer that resulted in feasible wing geometry, while still yielding maximum reduction 

in fuel burn prior to detailed vehicle analysis. Additional trade studies were performed to analyze landing 

and takeoff operations, and their respective effects on noise. With input from these trade studies, 

emissions, field length, mission fuel, and noise were used, in conjunction with the scenario weighting 

factors described in Section 4.1.2 and mission requirements, to calculate the SER for each vehicle. 

Vehicle characteristics were compared to both historical trends and other N+3 configurations. SER values 

were tabulated and sorted, and the vehicle with the highest SER was selected as the preferred N+3 

concept. 

Figure 6-1. Aircraft Vehicle Design Study Overview 

NASA/CR—2010-216798 65
6.1 Sizing and Performance Analysis 

6.1.1 Objective Function Effect on Aspect Ratio Sizing 

As described in Section 3.2, an objective function was needed to drive the design of configurations in 

the optimization process. Optimizing purely for mission fuel resulted in configurations with unattainably 

high aspect ratios, whereas optimizing for gross weight resulted in lower weight vehicles. Assuming that 

empty weight is a surrogate for vehicle cost, the issue becomes one of trading cost for fuel efficiency. An 

objective function was chosen as a weighted combination of both mission fuel and gross weight. This 

yielded aspect ratios that drove down induced drag, but still accounted for the empty weight of the 

vehicle. 

6.1.2 Advanced Liner Design 

A study was conducted into the potential of advanced inlet and fan duct liners to aid in the reduction 

of noise towards the N+3 goal levels. Using the methods previously described in Section 3.6, advanced 

inlet and fan duct liners developed for the International Aero Engines V2533-A5 were compared to 

current state-of-the-art liners. Knowledge gained from this study was used to estimate the acoustic 

benefits of implementing these liner designs into the N+3 configurations. Additional studies examined the 

possible weight penalties that would be associated with various liner types, and the impact on attenuation 

as a function of liner length. 

The baseline inlet liner was designed, and sound pressure reduction computed, using an eigenvalue 

expansion algorithm for the duct acoustic equation, combined with reference V2533 sound pressure levels 

at power levels corresponding to 0, 1,000, and 5,000 ft altitudes. The baseline inlet liner design was a 

2 DOF liner with a length of 15 inches and a septum inserted to about 1/6 of the core depth. 

Considerations were given to penalties associated with hard walls and splices associated with the 

manufacturing process. The advanced inlet liner took into consideration improved manufacturing 

processes that would eliminate these splices, combined with advanced materials which exhibit better 

absorptive properties. Due to projected advances in 

installation capabilities, the advanced liner is also 

capable of extending farther into the inlet, stretching 

up to 28 inches long. As a function of frequency, both 

baseline and advanced liner attenuations remained 

essentially independent of power setting. Their 

attenuations of the OASPL relative to a hard-walled 

inlet are shown in Table 6-1. 

A similar study explored liner design for fan duct applications, on both the inner and outer annulus of 

the duct. The baseline and advanced fan duct liners were both designed to a length of 50 inches, while the 

advanced liner was designed to achieve attenuation over a more broad range of frequencies. For the low 

altitude power setting, the attenuation of the liner as a function of frequency is shown in Figure 6-2. It can 

be seen that while the baseline liner achieves a higher peak attenuation, its SPL attenuation exceeds 8 dB 

only in the approximate range 1400 Hz – 4500 Hz. On the other hand, the advanced liner achieves this 

level of SPL attenuation over a broader range, 800 Hz – 5500 Hz. 

A summary of the OASPL reduction generated by the baseline and advanced fan duct liners relative 

to the hard wall case are shown in Table 6-2. Again, the reduction has been shown to be largely 

insensitive to power setting. 

Table 6-1. Baseline and Advanced Inlet Liner 

Overall Sound Pressure Levels Reduction

Power Setting Altitude, ft Hard Wall Baseline Advanced

0 --- -3.8 -6.9

1000 --- -3.3 -5.7

5000 --- -3.3 -5.8

OASPL, dB

NASA/CR—2010-216798 66
Figure 6-2. Baseline and Advanced Fan Duct Liner Sound Pressure Level Reduction 

A brief study of the weight impact of the liners 

showed that the baseline and advanced inlet liners 

would increase the weight of the engine by 0.33% and 

0.48%, respectively. The increased weight of the 

advanced inlet liner relative to the baseline liner is due 

largely to the increased length of the advanced liner. 

Assessing the weight impact of the fan duct liners, it 

was found that the baseline and advanced liners total 

up to 0.28% and 0.49%, respectively. Since the two 

liners have the same overall length, the weight difference stems from the additional thickness in the core 

of the advanced liner designed to target lower frequency ranges. 

Additional work focused on examining the effect 

of liner length on OASPL reduction for both the inlet 

and fan duct liners. The major conclusion drawn from 

this work was that increasing liner length results in a 

more dramatic reduction in the overall sound pressure 

level. However, it is noted that this increased length 

must be balanced against the increasing weight 

impact on the engine. Table 6-3 summarizes the 

OASPL reductions found in both the fan duct and 

inlet liners as a function of length, compared to the 

hard wall case. 

When more in-depth analysis is warranted, 

advanced studies can examine configurations in 

which the impedance of the liner varies along the 

length of the inlet or fan duct to target spatiallyvarying acoustic sources. 

6.1.3 Analysis of Engine Mounting Configurations 

6.1.3.1 Engine Mounting Configurations for Tube-and-Wing – The tube-and-wing 

configuration allows placement of the engines on the wing or the aft fuselage. Wing-mounting the engine 

is common for traditional aircraft of similar size and range. Appropriate tail geometry and volume 

0

2

4

6

8

10

12

14

16

18

100 1000 10000

Sound Pressure Level Attenuation, dB

Frequency, Hz

Baseline Liner

Advanced Liner

Table 6-2. Baseline and Advanced Fan Duct 

Liner Overall Sound Pressure Level 

Reduction 

Table 6-3. Impact of Liner Length On 

Overall Sound Pressure Level Reduction

Power Setting Altitude, ft Hard Wall Baseline Advanced

0 --- -5.2 -6.8

1000 --- -4.3 -5.7

5000 --- -4.3 -5.8

OASPL, dB

Length, in OASPL, dB Length, in OASPL, dB

15 -5.4 50 -5.6

20 -6.3 60 -6.0

28 -7.6 70 -6.3

30 -7.8 80 -6.7

35 -8.4 90 -7.0

40 -8.9 100 -7.2

45 -9.3

50 -9.7

Inlet Fan Duct

NASA/CR—2010-216798 67
coefficients were chosen for this configuration. To model fuselage-mounting accurately on the tube-andwing, a T-tail configuration was implemented in order to place the horizontal tail out of the engine 

exhaust. As a result, the taper ratio of the vertical tail was increased, while the root chord was held 

constant. This increase in volume and weight was needed to structurally support the raised horizontal tail. 

Engine placements and tail geometries were input into FLOPS to simulate the different engine mounting 

configurations. In addition to the resized tail weight, FLOPS internally computed a weight change for the 

wing and fuselage associated with the different engine mounting configurations. 

A trade study was conducted to quantify the performance and weight change associated with each 

engine mounting configuration. To accurately quantify this effect, a baseline tube-and-wing was used. In 

addition, both mounting configurations were analyzed with the addition of laminar flow. This would 

determine if the tail weight penalty associated with aft-mounting the engines could be overcome by its 

laminar flow benefits. 

The three-shaft turbofan was analyzed for four different design and technology combinations. As seen 

in Figure 6-3(b), wing-mounting affects the potential laminar flow region by creating a turbulent wake aft 

of the pylon. In this study, the size of the turbulent wedge is conservatively estimated based on nacelle 

diameter as opposed to pylon size. Aft-mounting the engines would allow for a greater extent of laminar 

flow on the wing, Figure 6-3(a). Depending on configuration, the removal of the engine could result in 

approximately an 8-10% increase in laminar flow on the wing. 

Figure 6-3. Wing Laminar Flow Extent for (a) Aft Fuselage Mounting, (b) Under-Wing Mounting 

As illustrated in Figure 6-4, given the tube-and-wing body with the three-shaft turbofan configuration, 

wing-mounting the engine resulted in the lowest fuel burned between configurations with no laminar 

flow, with a 50.9% reduction in mission fuel from the reference vehicle. Fuselage-mounting the threeshaft turbofan resulted in a 46.4% reduction in mission fuel from the reference vehicle. This difference in 

fuel is a direct result of the increased weight and drag associated with fuselage-mounting the engines. As 

discussed in Section 5.2.1, a T-tail configuration is necessary to move the horizontal tail out of the engine 

exhaust and subsequently increases the empennage structural weight and wetted area, leading to an 

increase in drag. 

When laminar flow is applied to both mounting configurations, wing-mounting is still preferred, and 

shows a 53.3% decrease in mission fuel compared to the reference vehicle where as fuselage mounting 

shows a 51.4% reduction. As mentioned above, the increased T-tail weight and drag of the fuselagemounted configuration results in a less fuel-efficient vehicle. Although this is offset somewhat by a 

greater extent of laminar flow on the wing, the net effect is that the wing-mounted configuration shows a 

1.9% fuel burn reduction improvement over the fuselage mounted configuration. 

(a) (b)

NASA/CR—2010-216798 68
Figure 6-4. Effect of Three-Shaft Turbofan Mounting Location on 

Tube and Wing Mission Fuel Burn With and Without Laminar Flow 

The open rotor was analyzed for three different configurations, as seen in Figure 6-5. However, the 

laminar flow field differs between the open rotor and three-shaft turbofan configurations. Specifically, 

laminar flow over the wing was assumed to be absent from the open rotor configuration. This is due to 

increased vibration in the wing and the disturbance of free stream air caused by the presence of rotors. 

Aft-mounting the engines, however, would allow for laminar flow on the wing, as seen in Figure 6-3(a). 

Figure 6-5. Effect of Open Rotor Mounting Location on Tube and Wing 

Mission Fuel Burn With and Without Laminar Flow 

A trade study was performed to quantify the fuel burn and weight change associated with changing 

the engine mounting location. These mounting configurations were compared to the baseline tube-andwing without laminar flow technology. In addition, analysis was performed to determine how much, if at 

all, full laminar flow over the wing offset the added structural weight associated with the aft-mounted 

configuration. 

As illustrated in Figure 6-5, given the tube-and-wing body with an open rotor, wing-mounting the 

engine results in the lowest fuel burn in the absence of laminar flow, reducing the fuel over the reference 

NASA/CR—2010-216798 69
vehicle 55.3%. Fuselage mounting the open rotor configuration in the absence of laminar flow resulted in 

a 51.4% fuel reduction compared to the reference vehicle. This slightly higher fuel burn reduction is a 

result of the increase in wetted area, drag, and weight associated with fuselage mounting the engines. 

Laminar flow was applied to the fuselage mounted configuration to see if the benefits associated with 

laminar flow would overcome the penalties associated with fuselage mounting. This configuration led to a 

53.6% reduction in fuel burn compared to the reference. Though the fuel burn is further reduced with the 

addition laminar flow, the wing mounting configuration still resulted in the largest fuel burn reduction. 

However, wing mounting the open rotor leads to significant complications. 

Due to the large rotor diameter, mounting the open rotor under the wing, which is common on tubeand-wing vehicles, would result in a rotor-tip ground clearance problem. Over-wing mounting would 

eliminate this problem, as engine is mounted on a pylon above the wing, increasing the distance between 

the ground and rotor tip. Additional benefits of over the wing mounting the engine would include a 

significant reduction in noise from noise shielding as mentioned earlier in Section 5.3.3.17 [Ref. 87]. 

However, cabin noise would increase significantly [Ref. 96]. Apart from an increase in cabin noise, 

mounting the engines above the wing would add complexity to servicing during engine maintenance and 

replacement. Structural complications would also arise, including unfavorable flutter characteristics 

[Ref. 97] and the need to mount the engine at a significant location above and aft on the wing in order to 

minimize detrimental aerodynamic interference [Ref. 98]. Though wing mounting the open rotor resulted 

in less fuel burn in this study, the complications associated with wing mounting this engine were seen to 

outweigh the improved fuel burn reduction and this configuration was not further investigated. 

6.1.3.2 Engine Mounting Configurations for Hybrid Wing Body – Because under-wing engine 

mounting is not practical with the hybrid wing-body, only a fuselage-mounted configuration was 

explored. The centerbody allowed for easy engine placement without having to reconfigure any aspect of 

the vehicle. There is no tail weight or size increase due to the HWB being a tailless vehicle. Additionally, 

fuselage-mounting provided improved noise shielding benefits over wing-mounting. Fuselage-mounting 

was determined to be the preferred configuration for both the open rotor and three-shaft turbofan engines. 

The centerbody of the HWB was assumed to have no laminar flow, due to the increased leading edge 

sweep angle and thickness-to-chord ratio. As a result, there were no trade studies performed between 

laminar flow and engine mounting. 

6.1.4 Laminar Flow Integration 

It was found that swept-wing laminar flow had a detrimental impact on the fuel burn of the HWB, 

whereas this technology led to a significant fuel burn improvement for the tube and wing vehicles. This is 

a result of the loss in takeoff and landing maximum lift coefficients associated with swept-wing laminar 

flow technology. Integrating laminar flow requires the removal of any upstream surface discontinuities 

that would otherwise trip the flow and create a zone of turbulence aft of this point. To ensure that this 

does not happen, the leading edge slat system needed to be removed. As a result of the leading edge slat 

removal, the lift generated by this high lift device was removed as well, penalizing the landing and 

takeoff maximum lift coefficients. Maximum lift coefficient is additionally penalized for using an airfoil 

more conducive to laminar flow. Wing area must be increased, with a subsequent penalty in empty 

weight, to meet the field length constraint with a reduced maximal lift coefficient. Consequently, the fuel 

required to carry out the mission will increase. Even with turbulent flow, a hybrid wing-body aircraft 

achieves a much lower maximum lift coefficient for takeoff and landing than a conventional tube-andwing aircraft. Therefore, the lift penalties associated with laminar flow are much more significant on a 

HWB than with the tube-and-wing configuration, as seen in Table 6-4. Mitigating these laminar flow 

penalties could result in reduced fuel burn for both configurations. Figure 6-6 illustrates that applying 

laminar flow to the hybrid wing-body with lift coefficient penalties results in a 13.55% increase in fuel 

burned. If the penalty associated with slat removal is mitigated through the use of wing morphing 

technology, steady leading edge blowing, or other forms of active or passive control, fuel burn is only 

increased by 4.32% from the reference configuration. Mitigating both laminar flow maximum lift 

NASA/CR—2010-216798 70
coefficient penalties actually results in a 6.8% 

improvement in the fuel burn of the hybrid 

wing-body compared to a fully-turbulent 

wing. Conversely, laminar flow yielded 

mission fuel reductions when integrated on the 

tube-and-wing vehicle while accounting for 

maximum lift coefficient penalties. As such, 

mitigating the laminar flow lift penalties on 

the tube-and-wing vehicle would lead to an 

increased fuel burn reduction. 

Figure 6-6. HWB With Laminar Flow Fuel Burn, Laminar Flow Penalty Mitigation Effects 

6.1.5 MDD and Sweep Analysis 

Drag divergence Mach number studies were performed for initial sizing purposes. A range of values 

for aspect ratio, sweep, lift coefficient, and weight were used to provide MDD sensitivities to wing sweep, 

thickness-to-chord ratios, and wing technology. Drag divergence Mach number calculations utilized the 

Korn Equation with 3-D effects and a defined range of parameters [Ref. 99]. Thickness-to-chord ratio was 

varied between 0.10 and 0.14, sweep was varied between 10 and 45 degrees, and lift coefficient was 

varied from 0.20 and 0.65. Both conventional and modern airfoils were examined. Conventional airfoils 

were comparable to the reference vehicle. Modern airfoil technology improves the drag divergence Mach 

number for constant sweep. 

A drag divergence Mach number for the tube and wing configuration was chosen to be 0.85. The 

selection of a drag divergence Mach number at this value allowed for the minimum cruise Mach number 

constraint of 0.75, discussed Section 4.2.3, to easily be met. A drag divergence Mach number of 0.85 

enables Mach 0.8 cruise capabilities. Through additional configuration and aerodynamic analysis, a 

thickness-to-chord ratio of 0.12 was decided upon. This ratio was in-line with historical values for tubeand-wing vehicles with similar size and passenger count. Interpolating the carpet plot illustrated in 

Table 6-4. Maximum Lift Coefficient Comparison 

With and Without Laminar Flow Penalties 

NASA/CR—2010-216798 71
Figure 6-7 yields a minimum sweep of 19.5 degrees when designing for a drag divergence Mach number 

of 0.85 and a thickness-to-chord ratio of 0.12. Further sweep definition was attained through refinement 

of the cruise lift coefficient for values consistent with historical values with similar vehicles. 

Figure 6-7. Initial Sizing Tube-and-Wing MDD Sensitivity Study 

The same drag divergence Mach number was 

chosen for HWB as the tube and wing 

configuration at 0.85 which would allow for 

Mach 0.8 cruise capabilities. Through additional 

configuration and aerodynamic analysis, a 

thickness-to-chord ratio of 0.12 was decided 

upon which is historically similar to NGC all 

wing vehicles. Figure 6-8 shows that for a MDD

of 0.85 and the given range of lift coefficients, 

the minimum allowable sweep was 25 degrees. A 

wing sweep was chosen to be 40 degrees which 

allowed for improved balance and increased 

stability and control. Figure 6-8 illustrates that a 

wing sweep of 40 degrees is more than sufficient 

for a drag divergence Mach number of 0.85. 

6.1.6 Landing and Takeoff Operations 

Detailed trajectories below 10,000 feet 

altitude were developed for each final vehicle 

configuration for takeoff and climb, as well as 

descent and landing. Detailed engine cycle 

characteristics were calculated at each 0.5 second increment and were used for predicting noise source 

levels for acoustic analysis. These trajectories were developed from 4-D linear interpolations of detailed 

FLOPS outputs. 

Takeoff and climb trajectories were calculated assuming a 100% power level for both engines. Part 

36 rules permit commercial aircraft to perform an engine cutback that would enable them to meet the 

minimum single-engine gradient in the event they lost an engine after the cutback. For descent and 

Figure 6-8. Initial Sizing HWB MDD

Sensitivity Study 

NASA/CR—2010-216798 72
landing, the vehicles initially followed a three degree glide slope in the approach configuration. At the 

outer marker; landing gear was extended, flaps were set to the landing setting, and thrust was adjusted as 

required to maintain a three degree descent at Vref + 10 KCAS. 

6.2 Analysis of Aircraft Emissions 

Nitrous oxide (NOx) emissions were computed for each aircraft and engine combination by 

examining the production of emissions during a simulated LTO cycle, defined by the International Civil 

Aviation Organization (ICAO), and described below. It was found that each configuration met the 

CAEP/6 NOx requirement, and that the advanced engine architectures far surpassed N+3 emissions goals 

due to their lean combustor technology, high overall pressure ratios, and high bypass capabilities. 

6.2.1 CAEP/6 NOx Requirement 

The Committee on Aviation Environmental Protection (CAEP) defined their CAEP/6 requirement on 

aviation emissions to take effect on engines entering service after 31 December 2007. Limiting the 

production of NOx per kN of SLS thrust, the requirement is a function of both overall pressure ratio and 

engine rated output at ISA sea-level static conditions. 

For engines with OPR < 30, the CAEP/6 requirement is defined by the following equation: 

Eqn. 29

Where Dp/F∞ is the NOx production in grams per kN of thrust, ∞ is the SLS OPR of the engine, and 

F∞ is the SLS rated thrust output. For engines with 30 < OPR < 82.6, the following CAEP/6 requirements 

apply: 

Eqn. 30

Additional restrictions are in place for OPR > 82.6; however, as none of the configurations 

considered in this study extend into this range, these restrictions will not be discussed. 

Shown graphically in Figure 6-9, the CAEP/6 requirement varies linearly on either side of OPR = 30, 

with the NOx requirement being less stringent with increasing OPR. In addition, the requirements for the 

higher thrust class engine are slightly more rigorous than those of the lower thrust-class, lower OPR 

range. 

Figure 6-9. CAEP/6 LTO NOx Requirements. 

F F kN F kN

F kN p FD

38.5486 .1 6823 .0 2453 .0 00308 ,26 7. 89 0.

16.72 .1 4080 , 89 0.

/

3 8

0

F F kN F kN

F kN p FD

46.1600 .1 4286 .0 5303 .0 00642 ,26 7. 89 0.

.1 04 0.2 , 89 0.

/

6 2

0

0

20

40

60

80

100

120

140

160

180

0 20 40 60 80 100

CAEP/6 NOx Requirement, g/kN

Overall Pressure Ratio

SLS Thrust = 16760 lbf

SLS Thrust = 24005 lbf

NASA/CR—2010-216798 73
6.2.2 Simulated LTO Cycle 

In computing the emissions performance of the 

configurations in question, the simulated LTO cycle defined by 

ICAO was used [Ref. 100]. This is a standard sea-level static 

engine cycle used to consistently quantify the emissions 

behavior of aircraft engines. Four cycle points are defined in 

terms of engine power setting, and time duration at each point, 

to simulate an entire LTO cycle (Table 6-5). The majority of the 

reference LTO cycle consists of maintaining the engine at a 7% 

idle condition, while only a small portion of time is used to 

simulate engine takeoff power. The emissions produced at each of these cycle points are summed for a 

total NOx production level, and divided by the rated SLS output of the engine to generate its NOx 

production rate. 

6.2.3 Emissions Results 

6.2.3.1 CFM56-3B1 – Emissions indices for the CFM56-3B1 at the LTO cycle points were taken 

from the ICAO engine emissions certification listed in the ICAO Engine Exhaust Emissions Data Bank 

[Ref. 101]. This report shows that the turbofan produces 17.7, 15.5, 8.3, and 3.9 g NOx per kg of fuel 

burned at the simulated takeoff, climb, approach, and idle conditions, respectively. In addition, the engine 

certification lists the fuel burn rate at each of these conditions: 0.946, 0.792, 0.290, and 0.114 kg/sec. 

Using these numbers, combined with the cycle point durations listed in Table 6-5, it is found that the 

CFM56-3B1 generates 3595 g NOx during the LTO cycle. Since this engine is rated at 20,000 lbf SLS 

thrust, its NOx production is 40.3 g/kN. With this rated thrust, and a SLS OPR of 22.44, the CAEP/6 NOx 

requirement for this engine is 48.33 g/kN. It can be seen that this particular engine is already certified 

below its CAEP/6 NOx requirement, by 16%. This shows that much progress must be made in combustor 

technology relative to this design to achieve the N+3 emissions goal. 

6.2.3.2 Open Rotor – NOx emissions index (EI), as a function of engine power setting, was 

computed for the lean combustor technology common to the open rotor and three-shaft turbofan engine 

cores based on research and past experience. Through staging of the combustor and the use of direct 

injection, an overall reduction in the EI at both low and high power settings leads to significant reductions 

in NOx production throughout the LTO cycle. In the range of thrust class for which the engine models 

were developed, fuel flow rate and thrust scale linearly, leading to the conclusion that the NOx production 

rate is independent of SLS thrust for each engine configuration. For example, an open rotor exhibiting a 

SLS thrust of 16760 lbf, with fuel flow rates of 0.331, 0.290, 0.149, and .093 kg/sec at the respective 

cycle points, leads to a NOx production of 10.63 g/kN. At SLS conditions, the OPR for this engine is 

35.18, giving this particular model a CAEP/6 NOx emission requirement of 73.72 g/kN. As a result, the 

engine outperforms the CAEP/6 NOx requirement by 85.58%, or 10.58% beyond the N+3 emissions goal. 

6.2.3.3 Three-Shaft Turbofan – Since the three-shaft turbofan shares a common core with the 

open rotor architecture, the NOx emissions indices at the LTO cycle points of interest remain the same as 

those described in the previous section. However, the fuel flow rates at the cycle points during SLS 

conditions are different, producing a NOx emissions level of 9.69 g/kN. At SLS conditions the OPR for 

this engine is 49.32. With a rated output of 14393 lbf SLS thrust, the engine CAEP/6 NOx emission 

requirement is 102.9 g/kN. As a result, the engine outperforms the CAEP/6 NOx requirement by 90.59%, 

or 15.59% beyond the N+3 emissions goal. For cases in which the engine is sized to a thrust class below 

the 89.0 kN inflection point in the CAEP/6 restrictions, the regulation would become less stringent, and 

the turbofan engine more easily surpass the N+3 emissions goal. 

Table 6-5. ICAO LTO 

Reference Cycle 

Power Setting Time, min

Takeoff 100% 0.7

Climb 85% 2.2

Approach 30% 4.0

Idle 7% 26.0

NASA/CR—2010-216798 74
6.3 Analysis of Aircraft Acoustics 

6.3.1 FAR Stage 4 Noise Requirement 

Limits imposed by the Federal Aviation Regulation (FAR) Part 36, Section 103 on the noise 

allowable by large transport category airplanes and jet airplanes are described in Appendix B to the rule 

[Ref. 102]. This regulation, currently broken into Stages 1-4, regulated the noise allowable by aircraft, 

depending on EIS date. The noise metric used to regulate commercial aircraft is Effective Perceived 

Noise Level (EPNL). The requirements are defined in terms of three component metrics: community, 

sideline, and approach exposure levels. Limitations on each of the EPNL components are enforced by the 

Federal Aviation Administration (FAA), and the values of these restrictions are a function of the number 

of engines propelling the aircraft and its gross weight. Since all aircraft in the current N+3 study carry two 

engines, the Stage 4 noise requirements, as a function of gross weight alone, are shown in Figure 6-10. 

Requirements formally dictate a maximum EPNL for each of the three component metrics, the Stage 3 

regulations of which are also plotted in Figure 6-10. To meet Stage 4 requirements, these component 

metrics must be met individually, as well as having a cumulative margin that exceeds 10 EPNdB below 

Stage 3. The N+3 goal, however, focuses only on cumulative noise, targeting a 71 EPNdB reduction over 

the Stage 4 requirement. 

Figure 6-10. FAR Part 36 Stage 4 Cumulative EPNL Noise Requirements 

6.3.2 Effective Perceived Noise Level (EPNL) Calculation 

As defined by FAR Part 36, Section 101, EPNL is derived from the maximum tone-corrected 

instantaneous perceived noise levels, documented in Appendix A2 of the standards. Using the tools 

described in Section 3.7, computations for EPNL are performed for the three conditions described above: 

community, sideline, and approach. For the approach condition, the measurement is taken during flyover 

at a point located 1.24 mi up range from the end of the runway. For the community measurement, the 

flyover noise is monitored at a point 4.04 mi down range from the point of brake release. During the 

sideline measurement, the noise is recorded at incremental locations 1476 ft on either side of the runway 

during takeoff. The cumulative EPNL is defined as the linear sum of these three components. 

6.3.3 Acoustics Results 

The following sections present the acoustic results for the reference, advanced tube-and-wing, and 

HWB vehicles. Sound pressure level (SPL) calculations of the different acoustic sources per vehicle 

allowed for the identification of the dominant sources for both takeoff-climb and descent-landing. Open 

rotor noise sources were unable to be computed given the insufficient and inadequate source models 

60

65

70

75

80

85

90

95

100

105

110

265

270

275

280

285

290

295

300

305

0 200000 400000 600000 800000 1000000

Stage 3 Component EPNL Requirement, 

EPNdB

Stage 4 Cumulative EPNL Requirement, 

EPNdB

Gross Weight, lbf

Cumulative

Community

Sideline

Approach

NASA/CR—2010-216798 75
available to date, so results described in this section correspond to vehicles with either the reference 

engine or the three-shaft turbofan installed. Jet and fan noise sources were the two most dominant 

component sources during takeoff and climb trajectories, due to the engines operating at maximum power 

and these sources being strong functions of thrust. Other acoustic sources (e.g., airframe noise sources) 

were not analyzed during takeoff and climb, as relative noise contributions from other vehicle 

components were negligible in comparison to jet and fan noise. Overall sound pressure level maps for the 

various acoustic sources during takeoff and approach are given in Appendix D. A summary of the 

acoustic analysis is shown in Table 6-6 and is discussed in further detail in the following subsections. 

Table 6-6. FAR Stage 4 EPNL Comparison 

6.3.3.1 Reference Vehicle – Acoustic analysis of the reference vehicle showed EPNL levels 

slightly higher than those available from open literature. The community, sideline, and approach EPNL 

values were 76.15, 87.15, and 106.39 EPNdB, respectively. However, the cumulative EPNL value of 

269.69 EPNdB does not account for the inclusion of an inlet acoustic liner, possibly explaining the 

slightly higher than expected value. 

Approach noise is found to be dominated primarily by landing gear noise, with a greater contribution 

from the main gear. The distinct forward-directed emission pattern of the acoustic signature of the landing 

gear are shown in Appendix D. The reference vehicle takeoff noise is dominated by both jet and fan noise 

sources for reasons mentioned in Section 6.3.3. 

6.3.3.2 Advanced Tube-and-Wing, Scaled CFM56-3B1, No Technologies – Acoustic analysis 

was performed for the advanced tube-and-wing configuration with a scaled CFM56-3B1 engine having 

no advanced technologies applied. Analysis on this vehicle, when compared to the analysis performed on 

the reference vehicle, allows for noise reduction to be quantified with respect to configuration change and 

re-sizing. The community, sideline, and approach EPNL for this configuration were 61.15, 87.15, and 

103.22 EPNdB respectively. The cumulative EPNL of 251.52 EPNdB does not account for the inclusion 

of an inlet acoustic liner. The increase in approach EPNL relative to the reference vehicle is a result of the 

trajectory of the re-sized vehicle. This is a function of the ATW’s geometry and weight. The landing gear 

noise levels are reduced from those of the reference vehicle as a result of the reduced gross weight of the 

advanced tube-and-wing. These comparisons can be seen in Appendix D. 

The most significant reduction in EPNL came from engine re-sizing. This is evident when comparing 

the jet and fan noise levels during takeoff and climb. The configuration change and re-sizing from the 

reference vehicle resulted in an overall reduction of 27.49 EPNdB. 

6.3.3.3 Hybrid Wing-Body, Scaled CFM56-3B1, No Technologies – Acoustic analysis was 

performed for the hybrid wing-body configuration with no advanced technologies and a scaled CFM56-

3B1 engine. Analysis on this vehicle, when compared to the analysis performed on the reference vehicle, 

allows for noise reduction to be quantified with respect to configuration change and re-sizing. The 

community, sideline, and approach EPNLs were 78.59, 101.08, and 95.27 EPNdB, respectively. The 

cumulative EPNL of 274.94 EPNdB does not account for the inclusion of an inlet acoustic liner. In 

addition, due to the exceptionally high fuel burn levels associated with the HWB, it will be shown that 

any noise benefits achieved through engine shielding would not drastically alter the overall SER for this 

vehicle relative to the other configurations. For this reason, complex noise shielding effects were not 

calculated. 

Metric (EPNdB) Reference Vehicle ATW No Technologies HWB No Technologies

Community 76.15 61.15 78.59

Sideline 87.15 87.15 101.08

Approach 106.39 103.22 95.27

Cumulative 269.69 251.52 274.94

NASA/CR—2010-216798 76
Increased gross weight and engine size led to increased far-field noise levels when compared to the 

ATW vehicle. Noise shielding offers potential noise reduction, but effects would largely be reflected in 

takeoff and climb noise analysis where engine noise is dominant. 

6.4 System Effectiveness Rating Results 

6.4.1 Technology Assessments 

In order to quantify the benefits of each technology at the system level, their individual effects were 

quantified with respect to the reference configuration on an SER basis. Using the definitions from Section 

3.8, the system effectiveness rating for a reference configuration implementing individual technologies 

were computed. Because of the complexity of the acoustic calculation, SEREPNL is assumed to be zero for 

all cases to be discussed below. Figure 6-11 shows the ratings of the individual non-propulsion 

technologies that were applied to the tube-and-wing vehicle. Sorted by increasing overall SER, the figure 

shows that M5 ultrahigh-performance fiber had the greatest impact on the overall scenario, followed by 

advanced metallic structural and subsystem alloys. Though swept-wing laminar flow results in an aircraft 

that burns less fuel than one using either aeroservoelastic structures or advanced metallics, the higher 

thrust requirement of the laminar flow aircraft drives a more stringent NOx requirement. Hence, the 

weighting of the emissions goal in the overall scenario actually drives the SER of the laminar flow 

aircraft lower than that of aircraft implementing these structural technologies. SERs pertaining to the 

Bright, Bold Tomorrow scenario are higher than the others due to the significant weighting of balanced 

field length, for which all aircraft are given full credit. The aerothermal inverted flow nozzle shows the 

least benefit, and is ranked below the no technology vehicle, due to the weight it adds to the propulsion 

system and the negative impact it has on thrust. However, the acoustic benefits of the technology are not 

reflected in this figure. Table 6-7 shows the SER values computed for each combination of technology 

and scenario. 

Figure 6-11. ATW Individual Technology Assessment 

0.00

0.10

0.20

0.30

0.40

0.50

0.60

System Effectiveness Rating

Overall King Carbon NIMBY Bright, Bold Tomorrow

NASA/CR—2010-216798 77
Table 6-7. ATW Technology SER Assessment, Without Noise 

Similarly, the non-propulsion-related technology effects have been measured on the SER basis 

relative to their effect on the HWB configuration. These results are presented in Figure 6-12. The same 

trends are noticed in these technologies, with the exception that the laminar flow case achieving a high 

SER assumes that all penalties associated with the integration of laminar flow have been mitigated, 

through flow control technology (for example, leading edge morphing, steady blowing, etc.). Laminar 

flow with the integration penalties assumed is ranked lowest among all technologies. In a more broad 

sense, it can be seen that the SERs calculated for the HWB configurations are notably smaller than the 

corresponding tube-and-wing cases. This is attributed to the benefits of accessing higher aspect ratio 

wings in driving toward the N+3 metrics compared to reconfiguring to a hybrid wing-body airframe. 

Table 6-8 shows the SER values computed for each combination of technology and scenario. 

Figure 6-12. HWB Individual Technology Assessment 

Overall King Carbon NIMBY Bright, Bold Tomorrow

Reference 0.0690 0.0724 0.0828 0.0517

ATW_CFM_AeroTherm 0.2624 0.1732 0.2172 0.3968

ATW_CFM_NoTech 0.2812 0.2050 0.2278 0.4109

ATW_CFM_Vortex 0.2820 0.2063 0.2283 0.4115

ATW_CFM_LGInt 0.2919 0.2215 0.2353 0.4189

ATW_CFM_CNT 0.2921 0.2233 0.2339 0.4191

ATW_CFM_3D 0.2926 0.2241 0.2343 0.4195

ATW_CFM_LGF 0.2947 0.2258 0.2373 0.4210

ATW_CFM_LIS 0.3026 0.2408 0.2399 0.4269

ATW_CFM_LamFlow 0.3092 0.2690 0.2267 0.4319

ATW_CFM_ASE 0.3202 0.2711 0.2493 0.4401

ATW_CFM_AdvMetal 0.3207 0.2715 0.2501 0.4405

ATW_CFM_M5 0.3595 0.3381 0.2708 0.4696

Advanced Tube & Wing

-0.10

0.00

0.10

0.20

0.30

0.40

0.50

0.60

System Effectiveness Rating

Overall King Carbon NIMBY Bright, Bold Tomorrow

NASA/CR—2010-216798 78
Table 6-8. HWB Technology SER Assessment, Without Noise 

6.4.2 Technology Suites 

Based on the results of the technology assessment, technology suites were developed to target noise 

and performance for each configuration. Each technology suite was developed with both the open rotor 

and three-shaft turbofan engine architectures. Technologies that were determined through the SER 

assessment to provide less of a system level benefit were discarded. For example, on a system level, the 

effect of landing gear component integration was determined to be less than that of landing gear fairings, 

and as a result, landing gear component integration is not carried through into the technology suites. 

These packages are shown in Table 6-9 for the hybrid wing-body, and Table 6-10 for the advanced tubeand-wing. 

Table 6-9. Major Technology Suites for Hybrid Wing-Body Configurations 

Overall King Carbon NIMBY Bright, Bold Tomorrow

Reference 0.0690 0.0724 0.0828 0.0517

HWB_CFM_LamFlow 0.1426 -0.0440 0.1649 0.3069

HWB_CFM_DEN 0.1706 0.0106 0.1733 0.3280

HWB_CFM_LamFlowNoSlat 0.1862 0.0410 0.1779 0.3396

HWB_CFM_CCW 0.1871 0.0427 0.1782 0.3403

HWB_CFM_Aerotherm 0.1899 0.0481 0.1790 0.3424

HWB_CFM_BLI 0.1992 0.0663 0.1818 0.3494

HWB_CFM_NoTech 0.2065 0.0807 0.1840 0.3549

HWB_CFM_Vortex 0.2074 0.0824 0.1843 0.3556

HWB_CFM_LGInt 0.2075 0.0825 0.1843 0.3556

HWB_CFM_3D 0.2153 0.0978 0.1868 0.3615

HWB_CFM_CNT 0.2154 0.0979 0.1867 0.3615

HWB_CFM_InletOpt 0.2154 0.0980 0.1867 0.3616

HWB_CFM_LGF 0.2172 0.0986 0.1901 0.3629

HWB_CFM_LIS 0.2262 0.1149 0.1939 0.3696

HWB_CFM_LamFlowNoSlatAF 0.2427 0.1477 0.1984 0.3821

HWB_CFM_ASE 0.2461 0.1463 0.2075 0.3846

HWB_CFM_AdvMetal 0.2536 0.1591 0.2114 0.3902

HWB_CFM_M5 0.2875 0.2146 0.2323 0.4156

Hybrid Wing-Body

Technology HWB No Tech

Scaled CFM56-3B1 X

Open Rotor X X

Three-Shaft Turbofan X X

Aeroservoelastic Structures X X

M5 Ultra-High Performance Fiber X X

Affordable Large Integrated Structures X X

3-D Woven and Stitched Composites X X

Advanced Metallics X X

Swept-Wing Laminar Flow

Carbon Nanotube Electrical Cables X X

Landing Gear Fairings X X X X

Landing Gear Integration

Aerothermal Concepts X X

Boundary Layer Ingestion X X

Steady Circulation Control

Inlet Optimization X X

Distributed Exhaust Nozzle X X

Vortex Generators X X

HWB Noise Tech HWB Performance Tech

CONFLICTS WITH LANDING GEAR FAIRINGS, LOWER SYSTEM EFFECT

NASA/CR—2010-216798 79
Table 6-10. Major Technology Suites for Tube-and-Wing Configurations 

6.4.3 Configuration Results 

Implementing the technology suites onto the two airframe configurations allowed for both the 

quantification of the technology benefits and a comparison of the overall system performance of the 

aircraft configurations. These comparisons are made in Figure 6-13, with corresponding data listed in 

Table 6-11. It was difficult to accurately determine the effect of an open rotor configuration on the overall 

aircraft noise, because open rotor noise prediction methods are still immature. Without the benefit of 

noise prediction for open rotor engines, it was possible only to compare configurations based on assumed 

EPNL differences between configurations using these engines. While not precise, this methodology did 

permit relative comparisons. Hence, all data described in this section is subject to the assumption that 

there are no EPNL differences between configurations when replacing a three-shaft turbofan with an open 

rotor. 

Figure 6-13. System Effectiveness Rating Comparison 

Between Candidate Configurations 

Technology ATW No Tech

Scaled CFM56-3B1 X

Open Rotor X X

Three-Shaft Turbofan X X

Aeroservoelastic Structures X X

M5 Ultra-High Performance Fiber X X

Affordable Large Integrated Structures X X

3-D Woven and Stitched Composites X X

Advanced Metallics X X

Swept-Wing Laminar Flow X X

Carbon Nanotube Electrical Cables X X

Landing Gear Fairings X X X X

Landing Gear Integration

Aerothermal Concepts X X

Boundary Layer Ingestion

Steady Circulation Control

Inlet Optimization

Distributed Exhaust Nozzle X X

Vortex Generators X X

ATW Performance Tech

NOT ANALYZED

PR ≈ 1 FOR PODDED INLETS

CONFLICTS WITH LANDING GEAR FAIRINGS, LOWER SYSTEM EFFECT

ATW Noise Tech

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.80

0.90

1.00

System Effectiveness Rating (SER)

Noise Fuel Burn

Emissions Field Length

NASA/CR—2010-216798 80
Table 6-11. Contributions to System Effectiveness Rating Calculations 

The most prominent result from this analysis is that the hybrid wing-body configurations are less 

effective on a system level with respect to the mission requirements. For example, with the preferred 

technology package, the advanced tube-and-wing realized an overall SER of 0.960 when using the threeshaft turbofan, while the HWB with the same engine only achieved an overall SER of 0.751. While both 

configurations achieved the N+3 goals of field length and NOx emissions, the ATW significantly 

outperformed the HWB in terms of fuel burn and noise. A comparison between engines showed that 

configurations using the open rotor marginally outperformed those implementing the three-shaft turbofan. 

The advanced tube-and-wing configuration with open rotors had an overall SER of 0.963, while the same 

airframe configuration powered by three-shaft turbofans had an SER of 0.960. 

Knowing that the system effectiveness ratings between the two most promising configurations 

differed by such a small amount, it was possible to calculate the threshold EPNL that would be required 

for the open rotor configuration to surpass the three-shaft configuration in terms of SER. For a range of 

EPNL differences between configurations (open rotor relative to three-shaft configurations), the resulting 

system effectiveness ratings are plotted in Figure 6-14, holding everything else constant. As long as the 

open rotor configuration is less than 2.65 EPNdB louder that the three-shaft turbofan configuration, the 

open rotor configuration edges out the three-shaft turbofan. However, for higher EPNL differences 

between configurations, the three-shaft turbofan is clearly the preferred engine. The general consensus 

among internal and external acoustics experts is that the open rotor will be significantly louder than the 

three-shaft engine, and that the actual acoustics of this engine will place it beyond the constraint imposed 

by the SER. For this reason, the open rotor architecture was not carried forward onto the preferred 

configuration. 

Fuel, lbf BFL, ft Stage 4 Cum EPNL, EPNdB CAEP/6 NOx, g/kN NOx, g/kN

Reference 25048 4999 274.6 269.7 48.3 40.3

ATW 3-Shaft Preferred 9409.7 4896 271.1 201.7 103.2 9.69

ATW 3-Shaft No Tech 12196 4996 272.3 251.5 100.9 9.69

ATW Open Rotor Preferred 9338.7 4999 272.0 201.7 77.6 10.61

ATW Open Rotor No Tech 12331.7 4998 273.4 251.5 73.6 10.61

HWB 3-Shaft Preferred 11818.5 4682 271.6 259.9 101.5 9.69

HWB 3-Shaft No Tech 15002.2 4831 273.1 266.9 97.6 9.69

HWB Open Rotor Preferred 11768 4989 272.6 259.9 73.3 10.61

HWB Open Rotor No Tech 14730 4745 275.1 266.9 69.3 10.61

Requirements and Metrics

Fuel BFL EPNL NOx SER, w/o Noise SER w/ Noise

Reference --- --- --- --- --- ---

ATW 3-Shaft Preferred 0.892 1.000 0.906 1.000 0.764 0.960

ATW 3-Shaft No Tech 0.733 1.000 0.292 1.000 0.711 0.769

ATW Open Rotor Preferred 0.896 1.000 0.919 1.000 0.765 0.963

ATW Open Rotor No Tech 0.725 1.000 0.309 1.000 0.708 0.770

HWB 3-Shaft Preferred 0.755 1.000 0.164 1.000 0.718 0.751

HWB 3-Shaft No Tech 0.573 1.000 0.087 1.000 0.658 0.675

HWB Open Rotor Preferred 0.757 1.000 0.179 1.000 0.719 0.755

HWB Open Rotor No Tech 0.588 1.000 0.116 1.000 0.663 0.686

SER Components Scenario Weighted SER

NASA/CR—2010-216798 81
Figure 6-14. Effect of Configuration ΔEPNL on System Effectiveness Rating 

6.5 Preferred Configuration 

A preferred configuration emerged from the conceptual design study that best meets the goals for the 

N+3 aircraft. Table 6-12 shows the requirements for the preferred configuration in terms of the N+3 

goals, and the achievements that were made towards each of these targets. The final configuration 

resulted in a fuel burn that was 63.49% below that of the reference vehicle, an EPNL that was 69.6 

EPNdB below the FAR Stage 4 requirement for aircraft of its weight class, NOx production that was 

90.59% below the CAEP/6 requirement for aircraft using engines of its OPR and thrust class, and a 

balanced field length of 4,999 feet, showing the potential of the aircraft to exploit metroplex operations. 

While not all N+3 goals were met with this configuration, tremendous strides were made towards these 

objectives while taking into consideration the reality of technology maturation, aircraft entry-into-service 

schedules, and the N+3 timeframe. 

Table 6-12. N+3 Goals Achievement 

6.5.1 Airframe Configuration 

The resulting airframe configuration from the N+3 

study was a tube-and-wing configuration optimized for 

its mission and using advanced propulsion systems, lightweight structural materials, and laminar flow technology 

(Figure 6-15). Designed for 120 passengers and a 1600 

nm mission, this advanced configuration uses high aspect 

ratio wings (AR = 12.7) made possible by lighter wing 

and nacelle structures. Additional technologies, including 

landing gear fairings, were used to mitigate the acoustic 

footprint of the aircraft during takeoff and landing. 

0.87

0.88

0.89

0.90

0.91

0.92

0.93

0.94

0.95

0.96

0.97

-40 -30 -20 -10 0 10 20 30 40

System Effectiveness Rating

EPNL Difference between Configurations, EPNdB

Open Rotor Noise + 

Performance Package

Three-Shaft Preferred 

Configuration Reference

Metric Overall Scenario Weighting Factor

Noise -71 EPNdB relative to Stage 4 200.3 EPNdB 201.7 EPNdB 69.6 EPNdB below Stage 4 20.00%

Fuel Burn 70% below reference vehicle 7514 lbf 9145 lbf 63.49% below reference vehicle 33.33%

Emissions 75% below CAEP/6 25.7 g/kN 9.69 g/kN 90.59% below CAEP/6 33.33%

Balanced Field Length exploit metroplex 5,000 ft 4,999 ft 13.33%

Requirement Actual Achievement

exploited metroplex

Figure 6-15. Preferred N+3 Configuration 

NASA/CR—2010-216798 82
6.5.2 Preferred Technology Package 

The preferred configuration is comprised of technologies that 

enable fuel burn reduction and vehicle noise source mitigation (see 

Table 6-13). Airframe technologies that reduce empty weight 

include large integrated structures, aeroservoelastic structures, M5 

ultrahigh-performance fiber, carbon nanotube electrical cables, 3-D 

woven and stitched composites, and advanced metallic alloys. 

Propulsive technologies include the three-shaft turbofan engine 

with compressor intercooling, CMC turbine blades, fuel-cooled 

cooling air, variable nozzle geometry, and lean combustor 

technology. Engine noise was reduced by both the variable 

geometry nozzle and advanced inlet acoustic liner, and total 

airframe noise was mitigated by landing gear fairings and the 

removal of leading edge slats attendant with swept-wing laminar 

flow. Laminar flow dramatically reduced fuel burn by reducing 

skin friction drag associated with turbulent flow regions. 

6.5.3 Weight Breakdown 

A summary of the weight for this aircraft is shown in Table 

6-14. It can be seen that because of the weight savings caused by 

the use of advanced materials, the structural weight comprises only 

23.9% of the takeoff gross weight. Besides the passengers and 

cargo, the largest contributor to the aircraft weight is systems and 

equipment, accounting for 24.0% of the total weight. The majority 

of this is due to furnishings and equipment, suggesting a possible target for future weight reduction. Table 

6-15 shows a breakdown of the aircraft mission weight. The maximum ramp weight for this aircraft is 

only 80,478 lbf, driven down both by the reduced mission fuel requirement brought about by engine and 

aerodynamic efficiency, and the reduced empty weight derived from significant use of advanced materials 

throughout the airframe. 

Table 6-13. N+3 Technology Suite 

Technology Suite

Three-Shaft Turbofan Engine

-Ultra-High Bypass Ratio of ~18

-CMC Turbine Blades

-Lean-Burn CMC Combustor

-Intercooled Compressor Stages

-Swept Fan Outlet Guide Vanes

-Fan Blade Sweep Design

-Lifhgtweight Fan/Fan Cowl

-Compressor Flow Control

-Active Compressor Clearance Control

-Variable Geometry Nozzles

Swept Wing Laminar Flow

Large Integrated Structures

Aeroservoelastic Structures

M5 Ultra High-Performance Fiber

Carbon Nanotube Electrical Cables

3-D Woven and Stitched Composites

Advanced Metallics

Landing Gear Fairings

Advanced Acoustic Inlet Liner

NASA/CR—2010-216798 83
Table 6-14. Weight Statement 

System/Components % TOGW Weight, lbf

Structural 23.9% 19264

Wing 7.9% 6365

Horizontal Tail 0.9% 691

Vertical Tail 0.4% 356

Fuselage 7.9% 6324

Landing Gear 4.4% 3552

Nacelle 2.5% 1976

Propulsion 6.3% 5058

Engines 5.5% 4453

Fuel Tanks & Plumbing 0.8% 605

Systems & Equipment 24.0% 19341

Surface Controls 1.7% 1377

Auxilary Power 0.8% 626

Instruments 0.7% 587

Hydraulics 4.4% 3581

Electrical 1.1% 865

Avionics 1.5% 1232

Furnishings & Equipment 11.8% 9518

Air Conditioning 1.7% 1357

Anti-icing 0.2% 198

Empty Weight 54.3% 43663

Flight Crew & Misc. 3.1% 2470

Crew & Baggage Flight, 2 0.6% 450

Crew & Baggage Cabin, 4 1.0% 800

Unusable Fuel 0.6% 492

Engine Oil 0.1% 70

Cargo Containers 0.8% 658

Operating Weight 57.3% 46133

Passengers & Cargo 31.3% 25200

Passengers 26.8% 21600

Passenger Baggage 4.5% 3600

Zero Fuel Weight 88.6% 71333

Fuel 11.4% 9145

Mission Fuel 11.4% 9145

Max Ramp Weight 100.0% 80478

NASA/CR—2010-216798 84
Table 6-15. Mission Weight Breakdown

6.5.4 Propulsion System 

The three-shaft turbofan engine design was selected as the propulsion system for the preferred aircraft 

configuration. Generating a maximum sea level static thrust of 14,393 lbf, the engine has a takeoff thrustto-weight ratio of nearly 6.47 when considering the weight of the engine system, and a takeoff specific 

fuel consumption of only 0.226 pph/lbf. At cruise the thrust-to-weight ratio is 0.91. The fan itself has a 

diameter of 59.5 inches, with 22 rotor blades and 54 stator vanes. Having a flange-to-flange length of 94.7 

inches, the total engine weight (including electrical and control systems) is 2226.7 lbf. With the nacelle, 

the overall length and diameter of the propulsion system are 164.7 inches and 76.7 inches, respectively. 

Heavy use of composites is assumed in the front-end fan module, while the LP compressor and HP 

compressor rely primarily on titanium alloys and inconel, respectively. The turbine section makes use of 

ceramic matrix composite blades, while the lean burn combustor is also constructed of CMC material. 

6.5.5 Configuration Comparisons 

Table 6-16 allows for the comparison between the calibration, reference, and preferred vehicle. 

Comparison between the 737-500 and the reference vehicle illustrates the resizing for reduction in 

passenger count and range in addition to meeting the mission goal. Comparison between the reference and 

preferred vehicle shows the effects of vehicle optimization on mission fuel and integration of advanced 

technologies including advanced propulsion. The wing reference area is shown to be drastically reduced 

while achieving a higher aspect ratio of 12.7. The ramp gross weight of the preferred vehicle is reduced 

by 39,692 pounds, owing to its reduced structural weight and reduced fuel requirements. 

Table 6-16. Tube-and-Wing Vehicle Comparison 

Mission Weight Breakdown 

Empty Weight 43663 lbf 

Operating Weight 46133 lbf 

Zero-Fuel Weight 71333 lbf 

Max Ramp Weight 80478 lbf 

Maximum Landing Weight 72430 lbf 

Units 737-500 Reference

Vehicle

Preferred 

Configuration

# Passengers [] 123 120 120

Range [nm] 2,400 1,600 1,600

Ramp Gross Weight [lb] 133,500 120,170 80,478

Empty Weight [lb] 68,860 67,350 43,660

Fuel Weight [lb] 42,186 25,048 9,144

Wing Reference Area [ft^2] 1,135 1,280 967.5

Wing Sweep [deg] 25 25 26

Wing Span [ft] 94.9 100.4 111.0

Wing AR [] 7.9 7.9 12.7

T/W Ratio [] 0.30 0.34 0.36

Max Wing Loading [psf] 117 94 83.2

Balanced Field Length [ft] 8,630 4,497 4,999

Landing Field Length [ft] 4,450 4,996 4,906

Fuel Optimization & Mission & Passengers Advanced Technologies

NASA/CR—2010-216798 85
6.5.6 Performance Summary 

6.5.6.1 Aircraft Mission Fuel Summary – The reference vehicle fueled for a 1600 nm range and a 

passenger count of 120 burned 25,048 pounds of fuel. NASA N+3 objectives set a 70% reduction in this 

fuel weight as the target mission fuel, or 7514.4 pounds. The mission fuel burn for the preferred 

configuration is 9145 lbf, reducing fuel burn 63.49% from the reference vehicle. 

A waterfall of mission fuel reduction is shown in Figure 6-16. The fuel burn progression is traced 

from the reference vehicle to the final preferred configuration. However, as mentioned in Section 5.3.3, 

technologies can be packaged in different combinations that change their relative benefit. The waterfall 

illustrated represents the consensus interpretation of the individual benefit of each technology. The threeshaft turbofan engine was the greatest contributor to fuel burn reduction with a 28.18% reduction in 

mission fuel. Wing resizing, reduced fuselage length, aspect ratio increase, and cruise Mach number 

reduction led to a fuel burn savings of 8.23%. Among the individual technologies, M5 ultrahighperformance fiber, swept-wing laminar flow, aeroservoelastic structures, and advanced metallic alloys 

were the greatest contributors to fuel burn reduction with 8.12%, 5.73%, 4.18%, and 3.40% reductions in 

mission fuel, respectively. A detailed segment fuel breakdown is shown in Table 6-17. 

Figure 6-16. Preferred Configuration Planform and Technology 

Contributions to Mission Fuel Burn Reduction 

Table 6-17. Fuel Burn by Mission Segment for Preferred Configuration 

0

5000

10000

15000

20000

25000

30000

Mission Fuel, lb

7514.4

-70%

-63.5%

N+3 GOAL

-8.23%

-2.23% -1.19% -1.15% -1.10%

-4.18%

-8.12%

-28.18%

-3.40%

-5.73%

25048 lb

NASA/CR—2010-216798 86
While not achieving the N+3 goal of 70% fuel burn reduction, tremendous progress was made in this 

direction. Though many candidate technologies were dismissed in the early stages of the study, their 

potential contributions to this goal are now evident. Their implementation on the preferred configuration 

may lead to the successful accomplishment of the N+3 mission fuel burn objective. Future analyses could 

explore additional key technologies and their integration. 

6.5.6.2 Aircraft Acoustic Summary – Analysis on the preferred configuration, when compared to 

that of the ATW vehicle described in Section 6.3.3.2, allows for the noise reduction associated with slat 

removal, advanced technologies, the three-shaft turbofan, and advanced inlet liner to be quantified. The 

acoustic inlet liner was scaled appropriately and applied to the three-shaft turbofan engine, resulting in 

significant fan noise reduction during takeoff and climb. In addition, the inlet liner reduced engine fan 

noise during descent and landing, but this effect was less significant during this stage of the flight. The 

inlet liner reduced the total vehicle EPNL by 6.0 EPNdB. Significant landing gear noise reduction was 

achieved due to the landing gear fairing integration, reducing total EPNL by approximately 6.2 EPNdB. 

The removal of leading edge slats further reduces cumulative EPNL by approximately 6.51 EPNdB. 

Furthermore, due to the reduction in gross weight and, consequently, required engine thrust, additional 

reductions in noise were achieved through lower exhaust jet velocities. 

The community, sideline, and approach EPNL values were 44.64, 67.65, and 89.46 EPNdB, 

respectively, resulting in a cumulative EPNL of 201.75 EPNdB. The preferred configuration reduces 

cumulative EPNL by67.94 EPNdB compared to the reference vehicle. The NASA N+3 acoustic metric, 

71 EPNdB below the Stage 4 requirement, is a function of vehicle gross weight. The Stage 4 requirement 

based on the gross weight of the preferred vehicle is 271.3 EPNdB, making the N+3 acoustic goal 200.3 

EPNdB. The preferred configuration generates an EPNL that is 69.55 EPNdB below its respective Stage 4 

requirement and is 1.45 EPNdB higher than the N+3 acoustic goal. A waterfall of acoustic reduction is 

shown in Figure 6-17 which traces the noise reduction from the reference vehicle to the final preferred 

configuration. In addition, a detailed comparison between the individual and cumulative acoustics metrics 

is shown in Table 6-18 for the preferred configuration and the vehicles discussed in Section 6.3.3. 

Figure 6-17. Preferred Configuration Platform and Technology 

Contributions to Community Noise Reduction 

Table 6-18. FAR Stage 4 EPNL Comparison 

Metric (EPNdB) Reference Vehicle ATW No Technologies HWB No Technologies Preferred Configuration

Community 76.15 61.15 78.59 44.64

Sideline 87.15 87.15 101.08 67.65

Approach 106.39 103.22 95.27 89.45

Cumulative 269.69 251.52 274.94 201.74

NASA/CR—2010-216798 87
6.5.6.3 Aircraft Emissions Summary – NASA N+3 objectives set a 75% reduction in NOx 

production below the CAEP/6 requirements as the target emissions goal. Section 6.2.1 discusses the 

calculation of the CAEP/6 requirements and vehicle emissions. For the gross thrust and OPR of the 

preferred vehicle, the CAEP/6 requirement is 102.92 g/kN, making the N+3 emission goal 25.73 g/kN. 

The preferred configuration generates NOx at a rate of 9.69g/kN, which is 90.58% below its 

respective CAEP/6 requirement and 62.33% below the N+3 emissions goal. This large reduction in 

emissions production is a result of the staged lean combustor. The CAEP/6 requirement and the N+3 

emissions goal are shown in Figure 6-18, along with the reference and preferred vehicles. 

Figure 6-18. CAEP/6 LTO NOx Emissions Requirement and N+3 Achievements 

6.5.6.4 Aircraft Field Length Summary – NASA N+3 goals call for the exploitation of 

metroplex operations. Based on the analysis described in Section 4.2, it was concluded that aircraft 

capable of using runway lengths of 5,000 feet would be optimum for operating in the metroplex flight 

environment. All vehicles were designed to meet the 5,000 foot balanced field length requirement and 

landing field length. Field length calculations assumed sea level and standard day conditions. The 

preferred configuration operates with a 4,999 foot balanced field length and a 4,906 foot landing field 

length, successfully meeting field length requirements. 

NASA N+3 Goal

75% reduction from CAEP/6

Reference

Preferred Configuration

NASA/CR—2010-216798 88
7 TECHNOLOGY MATURATION 

Key enabling technologies were identified in Section 6.5.2 that contributed toward the N+3 goals. In 

order for these technologies to be ready for EIS in the N+3 timeframe of 2030-2035, these technologies 

need to achieve TRL 6 by 2025. To ensure that these technologies achieve this TRL goal, top-level 

technology roadmaps were created that map out the major definitive milestones that must be met. The 

milestones include analysis, experiments, and scope of experiments planned out in high-level risk 

waterfalls to capture the critical path to technology insertion. Each of the maturation activities were 

assigned an appropriate duration to accomplish the task assuming reasonable levels of investment, and 

with the facilities available in the United States. It is envisioned that minor sideline activities will be 

performed alongside these major milestones to support the fundamental understanding of the physics 

involved, and to hedge the investment with other technologies that could offer a similar benefit. However, 

this study focused on the technologies that were found to offer the most probable success for the 

timeframe needed, for the mission and vehicle class, and with the most benefit to the configuration. These 

caveats should be kept in mind for evaluation of the TMPs developed here. 

7.1 Physical Landing Gear Fairings 

Physical landing gear fairings have been investigated in numerous recent evaluations at TRL levels of 

7 or 8, which is beyond the goal level for 2025. These fairings have been implemented mainly around the 

wheel axes in flight tests of full-scale aircraft intended for high design maturity. Further design efforts to 

improve these fairings should improve to lower the wake strength and turbulent kinetic energy distributed 

throughout the wake of landing gear. This activity should be coordinated with both virtual landing gear 

fairing efforts and component integration efforts. 

NASA/CR—2010-216798 89
7.2 Virtual Landing Gear Fairings 

Virtual fairings are estimated to be at a current TRL of 2, as low-level development has occurred only 

in controlled lab settings, on highly simplified models, and without realistic environmental affects (Figure 

7-1). Plasma actuators are viewed as a prime candidate for implementation since although they offer 

relatively low momentum addition compared to other active flow actuators, they may greatly affect the 

coherent structure of the shed vorticity so that the dipole air-structure interaction noise sources will be 

reduced. However, at this low TRL, it is viewed as a prime candidate in a series of actuator trials, so that 

the most effective control scheme can be developed. In combination with physical fairings, robust control 

may provide sufficient forcing to accrue significant tonal noise reduction in the several main components 

of gear noise, stemming from the wheels (low-frequency), struts (mid-frequency), and non-integral 

components (high-frequency). A challenge to this choice for active flow control application will be 

integrating an actuator that consumes little power, and creates less noise addition than the broadband 

sources it will attenuate. These, along with proving an essential control scheme to minimize weight and 

cost, are seen as the primary activities of the TRL 3 activity, in a laboratory bench-test setting. 

Figure 7-1. TMP for Virtual Landing Gear Fairings 

NASA/CR—2010-216798 90
7.3 Swept-Wing Laminar Flow Slat Integration 

Elimination of the slat by integrating virtual or seamless leading edge concepts is estimated at a 

current TRL of 2. This technology item spawned from a more detailed consideration of the laminar flow 

requirements, not initially assumed in the technology QFD study. Continued development of new 

approaches through experimental and computational approaches is envisioned, working towards 

approaches that are both effective at increasing CL on a laminar flow wing, and have a tolerable failure 

mode that prohibits immediate stall. Both passive and active techniques should be investigated, including 

virtual aerodynamic flow control, as well as actuated deployable geometry that is amenable to transition 

concerns on a relatively sharper leading edge typical of laminar airfoil designs (Figure 7-2). 

Figure 7-2. TMP for Swept-Wing Laminar Flow Slat Integration 

To add value at this level, and progress to TRL 3, a NASA-sponsored development task should 

include dissemination of a particularly promising wing/airfoil geometry so that university and other 

researchers can implement CFD and experimental studies across a common platform, and compare 

physical mechanisms and performance on a more equivalent basis. This task is expected to last several 

years, defining promising concepts that tradeoff performance with robustness and graceful failure modes. 

Modeling (CFD and empirical) efforts should be coordinated for inclusion on further system-level 

assessments. For TRL 4 maturation, several candidates should be scaled up for inclusion on a swept-wing 

laminar flow section test article, that demonstrates slatless augmented-lift capability in a moderately 

realistic integration at or approaching the full-scale takeoff and landing Reynolds numbers. A complete 

test matrix should be pursued that includes not only design conditions, but simulation of random failure 

modes, and the ability of the control system to augment failure. The outcome of this activity should verify 

that a realistic standalone integration can prove the estimated benefit and system-level cost (weight, 

complexity) with a design that attempts to be compatible with laminar flow without extensive 

documentation of this facet. Compatibility with laminar flow design will be explicitly confirmed in a 

TRL5 activity at a full-scale Reynolds number after validation and further development of candidate 

NLF/HLFC designs pursued in separate development activities. Therefore, transition measurements 

should be made at this point in an adequate ground-based facility, and/or coordinated leading edge 

NASA/CR—2010-216798 91
excrescence effects on transition location should be available at the appropriate Reynolds number. 

Control schemes should be validated and optimized at this level. The TRL 6 maturation activity should be 

a flight-test validation of a partial-section wing-glove integration. The primary goal is to confirm scaled 

model behavior and CFD predictions of the slat CLmax benefits, and the effects on laminar flow extent. 

These full-scale results will provide a crucial update to benefit models, and provide weight and 

integration effects, both of which should be implemented in system-level effects studies. 

7.4 Swept-Wing Laminar Flow Control of Crossflow Transition 

Crossflow transition delay is established as a primary hurdle to enabling swept-wing laminar flow. 

This technology was further defined after initial consideration of laminar flow concepts in the technology 

QFD study (Figure 7-3). 

Figure 7-3. TMP for Swept-Wing Laminar Flow Control 

HLFC using passive and/or active elements is approximated at a current TRL of 3, due to the dearth 

of realistic systems that will not greatly impose unacceptable system-level penalties of weight and suction 

or power levels. Advancement activities at this level to both delay transition while being capable of 

integrating high-lift techniques are required for downselecting capable candidate concepts designed for 

minimal vehicle-level system penalties. These activities should conclude in several years after 

considering several alternatives, and improving the understanding of the delay mechanisms and instability 

growth rate predictions. The culmination should be a set of designs that may be validated in a TRL 4 

subsystem validation that demonstrates control effectiveness at relevant simulated flight conditions. 

Facilities with low-turbulence qualities may need to be improved to offer sufficient freestream 

disturbance levels and character at the required Reynolds numbers for sufficiently confident validation. 

Crossflow transition delay models as a function of the major control parameters are key outputs of this 

demonstration. For advancement to TRL 5 maturity, a full-scale Reynolds number experiment of a swept 

wing section with an advanced design should be used to assess crossflow delay effectiveness. 

Implementation of a high-lift system (surrogate or functional) to verify integration and influence on 

transition mechanisms would be a highly desirable element at this level. TRL 6 attainment should be a 

product of a full-scale flight test demonstration, integrating results of key correlations developed in 

NASA/CR—2010-216798 92
previous activities, using a wing glove and an integrated high-lift system. This activity may be synergistic 

with the leading-edge slat elimination activity, in which a precursor system without high-lift integration 

may be tested as a baseline system for this activity, then leveraged for the further virtual slat integration 

and test. 

7.5 3D Stitched/Woven Composites 

Continued fabrication technique development of a stitched/woven composite structure will take this 

technology to a TRL 4 level in the very near term (Figure 7-4). A set of several candidate structures 

should be assessed at the NASA COLTS facility, potentially as a common activity with the large 

integrated structures technology development path. Wing boxes and main fuselage structures should be 

designed to a specific purpose (vehicle type and class), with competing candidate approaches used to 

verify the importance and magnitude of industrial knowledge and successful application of previous 

development guidelines, to reach TRL 5. A TRL 6 activity should be the successful instrumented 

implementation of successful candidates into a flight platform, with robust instrumentation, or perhaps in 

concert with novel health management and prognostic systems in place to gain valuable in-flight 

modeling static and dynamic characteristics for these novel structures that will ensure acceptable 

commercial acceptance. 

Figure 7-4. TMP for Stitched/Woven Composite Structures 

NASA/CR—2010-216798 93
7.6 Advanced Structural and Subsystems Metal Alloys 

At a current TRL of approximately 4 (depending on which material class of several), optimizations of 

alloy chemistries for structural and subsystem components should continue to investigate good candidates 

for limited sample production throughout the next several years (Figure 7-5). Near-term goals should 

focus on higher strength in compressive and tensile modes, as well as at elevated temperatures and at 

room temperatures. High-pressure, dynamic, hydraulic applications should be targeted as well. 2XXX and 

6XXX series alloys may offer some of these advances, while titanium-aluminum alloys, typical of 

structural use, may be investigated for use in subsystem applications such as hydraulics. Advancement to 

TRL 5 will be performed through advancing the selected chemistries through further thermo-mechanical 

optimization, and documentation of more thorough mechanical properties. TRL 6 will be reached through 

demonstrating these samples in limited, realistic loading and thermal environments, simulating the 

pressures and temperatures reached in intended applications throughout the airframe. Limited properties 

databases should be developed for comparison purposes to existing commodities. 

Figure 7-5. TMP for Advanced Structural and Subsystem Alloys 

NASA/CR—2010-216798 94
7.7 SMA Variable Geometry Nozzles 

Variable geometry nozzles using SMA materials to actuate a high-bypass ratio fan nozzle are 

estimated to be at a TRL of 4 (Figure 7-6). SMA characterizations should be performed to reliably 

(minimizing hysteresis) actuate at least a two-position nozzle practical for takeoff and cruise operations. 

This should be demonstrated in a full-scale exhaust system cold-flow sector test in a highly instrumented 

laboratory setting to confirm nozzle performance. Investigating the tradeoff between continuously varying 

nozzle area, and two or multiposition fixed positions should be performed by exploring multiple 

configurations over a 2 to 3 year period. CFD studies should be validated across the range of test 

conditions. An optimized closed-loop control system to manage SMA deflections should be demonstrated 

at this stage. This stage should be concluded with a warm-flow sector test to confirm performance in 

relevant, but not full-scale exhaust flow conditions. TRL 6 achievement should be attained by 

demonstrating an integrated variable area nozzle concept on a demonstration engine ground test. Thrust 

stand performance measurements and control system operation and tuning should be the primary 

verification metrics. Limited acoustic measurements should be made to confirm noise reduction trends. 

Figure 7-6. TMP for SMA Variable Geometry Nozzles 

NASA/CR—2010-216798 95
7.8 M5 Organic Fiber Composites 

The fabrication of composite layup panels using the high-performance M5 organic polymer fiber is 

currently estimated at a TRL of 3 (Figure 7-7). Further work to verify performance for aircraft panel 

applications should focus on using the status M5 fibers, which have known fabrication quality limitations, 

to develop layup processes with appropriate resins. Due to the excellent compressive properties, 

composite structural components should also be investigated at this point in system benefit studies. Fiber 

manufacturing development must occur in parallel to improve fiber performance and manufacturability to 

reinforce the TRL 3 level risk. Coupon testing to examine basic composite properties through break tests 

and surface examination, along with basic stress/strain testing should occur at this stage. For moving to 

TRL 4, improved fibers must be available starting in several years following coupon testing to advance a 

simple composite panel structure through compression and break testing, as well as verifying layup 

procedures and the expected resin-fiber interface properties. Examining the resilience to moisture and 

solvents should be preliminarily investigated, along with further surface examinations and sample 

preparations to gain confidence in suitability to finish applications. Advancement to TRL 5 will require 

building of a moderately complex array of structural samples to mimic various layups appropriate for 

different aircraft applications. More extensive properties testing will occur, along with failure mode and 

vibration examination for basic panel geometries. Fiber properties should have essentially reached “goal”

values at this point, so that the focus will be on successful integration. TRL 6 will pursue more complex 

panel development emulating scaled realistic implementations, alongside a structural allowables database 

development activity that will last at least several years. 

Figure 7-7. TMP for M5 Organic Fiber Composites 

NASA/CR—2010-216798 96
7.9 Carbon Nanotube Electrical Cables 

Carbon nanotube cables are currently estimated at a TRL of 3, and must be pursued in terms of both 

core wire performance, and outer mesh conductor performance (coaxial arrangement) (Figure 7-8). To 

reach TRL 4 for both of these components, at least two years of development is foreseen to assess 

feasibility of optimized designs for combined prototypes. Separate development will occur on each to 

adequately lower risk levels initially, and identify and isolate technical hurdles. Close collaboration 

between development activities should occur, and culminate in the demonstration of at least one but 

possibly several candidate wire prototypes for adequate standalone evaluation of properties. For acquiring 

a TRL of 5, performance of a subsystem prototype installation for a moderately complex electrical system 

should be evaluated. Reaching TRL 6 should address any deficiencies in the properties of the TRL 5 

through design and/or insulator optimization, and should culminate in the demonstration of a complete 

ground-based HITL subsystem. 

Figure 7-8. TMP for Carbon Nanotube Electrical Cables 

NASA/CR—2010-216798 97
7.10 Computational Inlet Optimization 

Early-design cycle optimization of inlet installations is estimated to be at a current TRL of 3 (Figure 

7-9). Researchers have demonstrated some components of low- and moderate-fidelity transport of fanface distortion through to the bypass duct and core compressor, but this has not been in concert with other 

activities required to bring this capability to TRL 4. This would also require adding additional multifidelity models of inlet losses and integration effects to estimate performance quickly for MDO 

applications. Further, the handling of complex geometry requirements (to handle BLI, podded, embedded 

configurations) passed automatically through an application program interface (API) that must be 

devolved into a set of modeling constraints should be a cornerstone component. To move to TRL 5, the 

capability of the multifidelity tool should include tools that are essentially used currently for preliminary 

design, for use in the concept evaluation process. This could be enabled by adapting to new highlyparallelized GPUs and nodal numerical PDE solution methods that offer the potential for more than an 

order of magnitude increase in CFD solutions appropriate for this moderate level of fidelity. This matured 

capability, accompanied with robust geometry handling and aero effects modeling including spill drag 

and afterbody effects, will advance the tool to TRL 6. Integration of enabling technology models for 

secondary effect, including flow control for serpentine inlet applications, should be performed at this level 

when the tool is mature enough to have a validation background and a history of successful trade studies. 

This TRL 6 task is considered ongoing with continuous improvements to fidelity as computational 

resources improve, and as experimental databases come online for novel integration concepts. 

Figure 7-9. TMP for Computational Inlet Optimization 

NASA/CR—2010-216798 98
7.11 Large Integrated Structures 

The current TRL of large integrated structures is estimated at 4, updated from an initial estimate used 

in the QFD study, which assumed TRL 3 (Figure 7-10). Continued development to further reduce risk at 

this level should include prototype development activities including reliability based design (RBD) 

methods for joint designs. Advanced structural models should be developed to predict performance at 

joint component and integrated structure levels. Weight models should be updated for system-level 

evaluations. A follow-on task for advancement to TRL 5 is the implementation of previous design and 

prediction tools to design/fab/test a large substructure in the NASA COLTS facility, in concert with the 

stitched/woven composites development activities. Validation of structural design through NDE and 

large-scale test results should provide adequate feedback for improving predictive analysis. Machining 

improvements should be pursued at this stage to meet TRL 6 requirements. A full-scale flight 

demonstration on a structural testbed aircraft, using improved prototype processes and performance 

predictions should advance this technology to TRL 6. 

Figure 7-10. TMP for Large Integrated Structures 

NASA/CR—2010-216798 99
7.12 Aeroservoelastic Structures 

The current TRL of commercially viable aeroservoelastic structures is estimated at 3 (Figure 7-11). 

Previous development of advanced state-space simulations, analysis methods, and flutter control law 

development should be leveraged to develop a simplified high aspect ratio wing design to be evaluated by 

testing of a cantilevered semispan wind-tunnel model. A static aeroelastic scaling of this subscale model 

should be sufficient for this activity. Moderate drag reduction and load control goals, with active or 

simulated sets of control surface deflections and twist, should be pursued. Comparison should be made to 

the best passive aeroelastic design candidate, in a similar lead-up testing effort, to quantify the risk and 

benefits of the active implementation over the passive tailored aeroelastic wing. Parallel refinement of 

requirements for an integrated system should be performed and incorporated into future model designs to 

a practical extent. By demonstrating adequate modeling and benefits, TRL 5 will be earned through a 

further design sophistication level, including a symmetry mounting condition with active and dynamic 

demonstration of load and drag effects, and integration into a status N+3 design candidate. Subsystem 

integration and laminar flow design integration should be demonstrated at a high level. Continued design 

tools development and increasing control-law study will culminate in reaching TRL 6 with a flight test 

article at realistic flight Reynolds numbers in a subset of the intended flight envelope. 

Figure 7-11. TMP for Aeroservoelastic Structures 

NASA/CR—2010-216798 100
7.13 Lightweight Fan Structure and Cowl 

The design maturity of the lightweight fan structure and cowl is estimated at a current TRL of 3 

(Figure 7-12). Preliminary design and analysis, including impacts on the isolated engine system in terms 

of weight, flowpath losses, external aerodynamic effects, and integration challenges, should be performed 

to raise the TRL to 4. CFD databases should confirm the design for a subset of driving parameters, and 

possibly include DOE to develop a response surface for interdisciplinary modeling. Some component 

development/modification to verify design assumptions should be performed for key elements, likely the 

AGB and fan cowl. TRL 5 should be reached by component demonstration on an available engine 

simulator with a reference nacelle design, indicating compatibility with an available gas generator, and 

proving predictions and limited integration challenges. Status goals should be achieved for the weights 

and flows at the design conditions, and at a limited set of off-design points. Advancement to TRL 6 

should be made through a concerted design effort to establish the validity of an advanced prototype for 

performance and acoustic goals. Pylon integration should be investigated with a reference pylon, for 

impact on scrubbing losses, and aeroacoustics. A simulated full-scale engine test at static conditions will 

be the definitive ground performance assessment activity. Acoustic microphone measurements alongside 

plume diagnostics, along with the incorporation of a thrust stand will confirm performance predictions. 

Figure 7-12. TMP for Lightweight Fan Structure and Cowl 

NASA/CR—2010-216798 101
7.14 Intercooled Compressor Stage 

Compressor intercooling for high-OPR engines is estimated to be at a current TRL of 3 (Figure 7-13). 

Evaluation of subscale concepts to investigate pressure loss and cooling performance should be 

performed in the next several years. This should occur in simulated design flow (flow rates, pressure) 

conditions downstream of HPC installation to bring to TRL 4. CFD simulations should be carried out to 

understand pressure loss mechanisms and improve heat transfer rates. TRL 5 should be attained by 

installing a representative prototype intercooler into a temperature and pressure simulator rig, at large 

scale, to identify performance map across engine flow conditions of interest. Design should pursue 

appropriate goal weights. TRL 6 attainment should occur by a full-scale rig implementation in a high 

OPR (>50) core engine compressor simulator. 

Figure 7-13. TMP for Intercooled Compressor Stage 

NASA/CR—2010-216798 102
7.15 Cooled Cooling Air 

Cooled cooling air is currently estimated at a TRL of 5 (Figure 7-14). To advance to a TRL of 6, 

demonstration on an integrated, ground-based engine should be performed. Assessment of flow levels and 

heat exchanger pressure losses, with surrogate cooling fluids representative of cracked fuels (if none are 

available) should be performed. A low-loss, lightweight heat exchanger design should be validated as a 

key outcome. 

Figure 7-14. TMP for Cooled Cooling Air 

NASA/CR—2010-216798 103
7.16 Endothermic Fuel Cracking 

The use of endothermic jet engine fuels as a physical heat-sink through a decomposition reaction to 

absorb sensible heat (“cracking”) is estimated at a current TRL of 3 (Figure 7-15). Developing an array of 

potential candidates of jet fuels in bench-test burners measuring heat sink capability and exothermic 

properties would elevate this technology to a TRL of 4. Simultaneous efforts to deoxidize the fuel and 

pursue other coke-mitigation strategies, although not detailed here, should be made to sufficiently reduce 

coke deposition due to the oxidation mechanism across the temperature range relevant to an inline 

turbofan fuel heat exchanger for cooled cooling air. To achieve TRL 5, demonstrating lean-burning fuels 

in representative conditions of critical design points (takeoff, cruise) in the laboratory setting would be 

required, while maintaining acceptable increases in heat-sink capabilities over conventional hydrocarbon 

jet fuels. Complete emission studies should be performed to characterize the reactions and provide 

detailed data for kinetic reaction models. To reach TRL 6, application to an advanced burner 

configuration, confirming both low NOx production and high heat-sink capability must be demonstrated, 

while simultaneously confirming low coking. 

Figure 7-15. TMP for Endothermic Fuel Cracking 

NASA/CR—2010-216798 104
7.17 Lean-Burn Staged Combustion Liner 

This technology item, estimated at a current TRL of 5, is a technology synthesis of staged combustion 

(TRL 6) as well as the CMC burner liner (TRL 5) required to allow operation of the staged combustor at 

cruise conditions in a well-mixed injection state to promote lower NOx. Thus development of the CMC 

liner technology should occur in parallel with this activity so that the design for staged combustion, which 

will allow huge NOx reductions across the power setting range, can assume CMC availability (Figure 

7-16). To raise the staged combustion design to TRL 6, several steps should occur to validate the 

components in an increasing level of combined system complexity, eventually culminating in a 

demonstration involving lean direct-injection, staged combustors in an optimized geometric configuration 

and count, and using a CMC burner prototype. The first activity should be the design and test of injector 

concepts for use in staged combustion on a flame tube to validate the design and also investigate the 

compatibility with alternative fuels. This level of component test is deemed a higher TRL than similar 

N+3 TMP activities due to the extensive investigation performed under the DOE Advanced Turbine 

Systems program . The next development activity should use the diagnostics data from the previous study 

to evaluate performance on a two-cup sector test at high OPR with focus on integration, ignition, lean 

blowout, and sub-idle efficiency. A CMC liner should be demonstrated at this phase in a follow-on 

activity prior to larger-scale or subsystem evaluations. Finally, TRL 6 will be reached by demonstrating a 

well-integrated design in a scale engine simulator near design OPR and at off-design conditions. A CMC 

liner integration should be included at this stage to verify the wall-cooling flow requirements and 

integration with the staged injectors. 

Figure 7-16. TMP for Staged Combustion 

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8 CONCLUSIONS AND RECOMMENDATIONS 

A conceptual commercial passenger transport study was performed to define a single vehicle capable 

of meeting the future demands of entering into market service in the 2030-35 timeframe. The primary 

vehicle requirements that drove technology selection, configuration selection, and optimization were 

derived from the published NASA N+3 goals. These goals were further refined based on context derived 

from an overarching future scenario that captured likely future economic and operational factors 

suggesting that fuel burn and NOx emissions would be primary market drivers in this timeframe. 

Simulations of the vehicle in the National Airspace with the implementation of NextGen air traffic 

management capabilities were used to provide additional requirements on vehicle sizing by investigating 

the impact of the metroplex operations model to projected traffic levels. Results of this study suggested 

that a 5,000 ft runway capability would be sufficient to leverage existing reliever airports and meet the 

demands of future air traffic. Furthermore, the inclusion of aircraft with reduced cruise speed would not 

hinder the predicted air traffic, while achieving better levels of specific fuel consumption. This paradigm 

shift towards a slower commercial transport at higher altitude is a key driver towards reaching the fuel 

burn objectives of the N+3 generation of aircraft. 

A single design that best fills a broad, primary market need for 2030-2035 entry-into-service was 

developed after considering various combinations of technology suites, revolutionary and traditional 

configurations, and requirements trades. This design is characterized by an advanced technology tube and 

wing configuration with a three-shaft advanced technology turbofan. The resulting aircraft achieves a 

64% reduction in fuel burn, a 70 EPNdb reduction in noise, and a 91% reduction in NOx when compared 

to a reference aircraft incorporating technologies representative of aircraft in service in 2010. Field 

performance for this preferred concept enables operation from metroplex airports. 

Technologies targeting engine performance and vehicle empty weight were most beneficial in 

reducing fuel consumption. Estimates for structural weight reductions in the N+3 timeframe led to 

configurations in which the takeoff gross weight is constrained by passenger count, equipment, and 

furnishings. This result suggests targets for future weight reduction strategies. While the primary 

scenario-weighted N+3 metrics were determined to be fuel burn and emissions, it must be noted that 

realistically cost is a major driver in aircraft design. While structural technologies can permit access to 

higher aspect ratios than have been seen in the past, the weight increases accompanying larger wings will 

also drive cost. Hence, the aircraft that burns the least fuel will, in all likelihood, not be the least 

expensive aircraft. 

Noise was found to be dominated by fan sources and landing gear. To mitigate these sources, two 

degree-of-freedom liners were applied to address fan noise, and landing gear fairings targeted the gear 

source. In order to take advantage of swept-wing laminar flow, leading edge slat technology had to be 

excluded at the expense of takeoff aerodynamics. However, the removal of slat noise drove down 

airframe noise appreciably. With the preferred configuration, approach noise seems to be the dominating 

contributor to overall EPNL. 

A large contributor to reducing fuel burn was the three-shaft advanced technology turbofan engine. 

Driving high propulsive efficiencies with its ultrahigh-bypass ratio and thermal efficiencies with its high 

overall-pressure ratio, the engine achieved specific fuel consumption levels that were fractions of those 

realizable in modern engine technology. This engine also employs a suite of advanced propulsion 

technologies, including CMC turbine blades, compressor intercooling, and active compressor clearance 

control. 

Ever more stringent emissions requirements have been addressed by the use of staged combustion 

technology. Through work at Rolls-Royce Liberty Works, in conjunction with NASA Glenn Research 

Center, this technology has been demonstrated to lead to significantly reduced NOx emissions. This is 

accomplished by optimizing the burner for lean combustion at multiple flight conditions, including low 

NASA/CR—2010-216798 106
power settings during which the bulk of LTO NOx is formed. In addition, this configuration allows for 

enhanced lean blowout and ignition characteristics. 

A promising technology to reduce fuel burn was the open rotor engine. However, when compared on 

the basis of a system effectiveness rating, it was determined that the open rotor would earn its way onto 

the aircraft only if it was marginally louder than the three-shaft turbofan. With limited experimental 

acoustic research available on the topic of open rotor noise, a conservative assumption was made that its 

noise characteristics will be louder than the three-shaft turbofan. However, this critical assumption 

requires validation, and it is recommended that further research be conducted in an attempt to quantify the 

noise associated with this type of propulsion system. Through computational fluid dynamic modeling and 

noise source quantification, acoustic propagation analysis can predict the cumulative noise associated 

with this type of engine. In this manner, a more definitive comparison can be made between the open 

rotor and three-shaft turbofan engines, and the configurations they drive. 

Significant reduction in empty weight was a key contributor to reduced fuel burn. The most 

promising structural technology is the M5 ultrahigh-performance fiber. Because of the potential empty 

weight reduction of the aircraft stemming from the implementation of this material, exploratory studies 

are recommended with the goal of developing sizeable panels. This type of feasibility study could go far 

towards realizing the empty weight savings possible through the development of advanced materials. 

Other high impact structural technologies included aeroservoelastic structures and advanced metallic 

alloys, whose development is expected to involve less commercial risk than M5. 

Trade studies showed that there are complications to achieving significant laminar flow. Maintaining 

a smooth leading edge conducive to laminar flow precludes the use of conventional leading edge devices 

for high lift. The removal of a leading edge slat and reshaping of the airfoil for laminar flow must be 

compensated with some degree of leading edge flow control or morphing structure to prevent the wing 

from having to be oversized to meet takeoff requirements. These complications appeared to be more 

challenging for a hybrid wing body configuration. Steady or unsteady leading edge stagnation point 

blowing has been shown in a few studies to result in improved maximum lift coefficient and stall angle. 

This technology is recommended for further experimental work to complement ongoing laminar flow 

research. 

Overall, this conceptual design study shows that significant progress is achievable for designing and 

building aircraft to meet the needs of the 2030-2035 time frame. Technology can be applied to make a 

more socially and environmentally responsible air transportation system that meets the needs of the 

future. Technology roadmaps show a logic path forward to develop the technologies to enable this level 

of achievement. Northrop Grumman looks forward to being a partner with NASA in the innovation, 

development, and demonstration of these critical technologies.

NASA/CR—2010-216798 107
Appendix A 

Airframe Selections 

NASA/CR—2010-216798 108
Appendix A Airframe Selections 

A.1 Airframe Technology Selections 

 Airframe Concept #1.01 Advanced Metallic Structural and Sub-System Alloys 

 Airframe Concept #1.02 Fiber Metal Laminates 

 Airframe Concept #1.03 Metallic Digital Direct Manufacturing 

 Airframe Concept #1.04 Structural Materials Laser Surface Treatment 

 Airframe Concept #1.05 3-D Woven/Stitched Composites 

 Airframe Concept #1.06 M5 Ultrahigh-Performance Fiber 

 Airframe Concept #1.07 Light-Weight Foamed Composites 

 Airframe Concept #1.08 Conformal Load-Bearing Antenna Structures 

 Airframe Concept #1.09 Integrated Aeroservoelastic Structures 

 Airframe Concept #1.10 Open-Celled and Stamped Composite Structures 

 Airframe Concept #1.11 Post-Buckled Composite Structures 

 Airframe Concept #1.12 Affordable Airframe Large Integrated Structures 

 Airframe Concept #1.13 Direct Manufacturing/Laser Sintering of Polymeric Materials 

 Airframe Concept #1.14 Carbon Nanotube (CNT) Electrical Cables 

A.2 Aerodynamic Technology Selections 

 Aero Concept #2.01 Steady Fluidic Separation Control 

Aero Concept #2.02 Synthetic Jet Separation Control 

 Aero Concept #2.03 Upper Surface Exhaust Nozzle and Flap 

Aero Concept #2.04 Fluidic Thrust Vectoring for Upper Surface Exhaust Nozzle and 

Flap 

Aero Concept #2.05 Lower Surface Exhaust Nozzle and Flap 

Aero Concept #2.06 Distributed Exhaust Nozzle and Flap 

Aero Concept #2.07 Steady Circulation Control 

Aero Concept #2.08 Unsteady Circulation Control 

Aero Concept #2.09 Swept-Wing Laminar Flow 

Aero Concept #2.10 Riblets and Microvortex Generators 

Aero Concept #2.11 Active Drag Reduction 

Aero Concept #2.12 Modeling for Inlet Optimization 

Aero Concept #2.13 Inlet Flow Control (Microvortex Generators, Jets, etc.) 

Aero Concept #2.14 Channel Wing 

NASA/CR—2010-216798 109
A.3 Noise Technology Selections 

Noise Concept #3.01 Distributed Exhaust Nozzle (DEN) 

Noise Concept #3.02 Shielding (Fan, Jet, other Propulsive Sources) 

Noise Concept #3.03 Chevrons 

Noise Concept #3.04 Ejectors 

Noise Concept #3.05 Fan Blade Sweep Design 

Noise Concept #3.06 Fan Blade Flow Control 

Noise Concept #3.07 Geared Turbine 

Noise Concept #3.08 Variable Geometry Nozzles 

Noise Concept #3.09 Porous Plug Designs 

Noise Concept #3.10 Aerothermodynamic Concepts (Inverted Flow Nozzle, 

Thermal-Acoustic Shield) 

Noise Concept #3.11 Ultra-High Bypass Turbofans 

Noise Concept #3.12 Acoustic Excitation 

Noise Concept #3.13 Curved, Scarfed Inlet Ducts 

Noise Concept #3.14 Moveable Winglets 

Noise Concept #3.15 Deployable Vortex Generators 

Noise Concept #3.16 Brush-type Trailing Edges 

Noise Concept #3.17 Trailing Edge Serrations 

Noise Concept #3.18 Porous Flap Side Edges 

Noise Concept #3.19 Flap Tip Fences 

Noise Concept #3.20 Slat Trailing Edge Serrations 

Noise Concept #3.21 Slat Cove Fillers 

Noise Concept #3.22 Slat Gap Liners 

Noise Concept #3.23 Porous Slat Pressure Surfaces 

Noise Concept #3.24 Slat Tip Fences 

Noise Concept #3.25 Continuous Mold-Line Linkages 

Noise Concept #3.26 Streamlined Slat Tracks 

Noise Concept #3.27 Micro-tab Features 

Noise Concept #3.28 Wing Morphing (Shape Memory Alloys, Piezoelectrics) 

Noise Concept #3.29 Active Noise Cancellation 

Noise Concept #3.30 Landing Gear Assembly Fairings 

Noise Concept #3.31 Landing Gear Assembly Component Integration 

Noise Concept #3.32 Wheel Caps 

Noise Concept #3.33 Wheel Gap Filler 

Noise Concept #3.34 Landing Gear Assembly Plasma Fairings 

NASA/CR—2010-216798 110
Noise Concept #3.35 Linear Acoustic Liner 

Noise Concept #3.36 Single Degree-of-Freedom Liner 

Noise Concept #3.37 Double Degree-of-Freedom Liner 

Noise Concept #3.38 Multiple Degree-of-Freedom Liner 

Noise Concept #3.39 Metal Foam Liner 

Noise Concept #3.40 Nanotube Liner 

A.4 Propulsion Technology Selections 

Prop Concept #4.01 Conformal Polymer Solar Cell Power Augmentation 

Prop Concept #4.02 Constant Volume Combustion/Wave Rotors 

Prop Concept #4.03 Plane Jet Combustion 

Prop Concept #4.04 Swept Fan Outlet Guide Vanes 

Prop Concept #4.05 Open Rotor with Counter-Rotating Fan Blade Rows 

Prop Concept #4.06 Electric Subsystems 

Prop Concept #4.07 Oil-Less Engine 

Prop Concept #4.08 Boundary Layer-Ingesting Inlets 

Prop Concept #4.09 Embedded/Conformal Inlets 

Prop Concept #4.10 Thrust Vectoring Nozzles 

Prop Concept #4.11 Thrust Vectoring Open Rotor 

Prop Concept #4.12 Alternative Hydrocarbon Biofuels 

Prop Concept #4.13 Compressor Flow Control 

Prop Concept #4.14 Third Engine Stream 

Prop Concept #4.15 Lightweight Fan Structure and Cowl 

Prop Concept #4.16 Low Tip Speed Counter-Rotating Ducted Fan 

Prop Concept #4.17 Four-Shaft Turbine 

Prop Concept #4.18 Active Compressor Clearance Control 

Prop Concept #4.19 Combined/Compound Cycles 

Prop Concept #4.20 Turbo-Electric Distributed Fans 

Prop Concept #4.21 Fuel Cell APU 

Prop Concept #4.22 Hybrid All-Electric Propulsion 

Prop Concept #4.23 TO/C-Assist Propulsion Unit 

Prop Concept #4.24 Variable Pitch Fan Blades 

Prop Concept #4.25 Cryogenic Motors 

Prop Concept #4.26 Intercooled Compressor Stages 

Prop Concept #4.27 Porous Ceramic Materials 

Prop Concept #4.28 Low-Noise Fan Blades 

NASA/CR—2010-216798 111
Prop Concept #4.29 Combined Steam Cycle 

Prop Concept #4.30 Ceramic Matrix Composite Turbine Blades 

Prop Concept #4.31 Water Injection 

Prop Concept #4.32 Staged Combustors 

Prop Concept #4.33 NonTurbine Distributed Fans 

Prop Concept #4.34 All Electric (no AGB) 

Prop Concept #4.35 Turboshaft-Powered Direct-Drive Distributed Fans 

NASA/CR—2010-216798 112
Appendix B 

Technology Dispositions 

NASA/CR—2010-216798 113
Appendix B Technology Dispositions 

B.1 Constant Volume Combustors/Wave Rotors 

Constant volume combustion (CVC) engines use the detonative mode of combustion to access higher 

thermal (and overall) engine efficiencies [Ref. 104]. Inherent unsteadiness in the detonation process 

makes it difficult to capitalize on the theoretical performance and emissions benefits associated with this 

type of engine. CVC engines will typically have a large number of independently controlled detonation 

tubes. Constant volume combustion/wave rotor engines could be incorporated as augmenters to current 

turbofan technologies, or as stand-alone propulsion systems. In either case, their integration into an 

aircraft will be extremely complex and system-dependent. Potential benefits include reduced fuel 

consumption, driven by increased thermal efficiency, and reduced emissions due to short residence times 

associated with the detonation process [Ref. 105]. Penalties include increased engine complexity, 

potential for increased noise due to unsteady shock waves, and structural considerations due to impulsive 

shock loading of the engine nozzle. 

Theoretical thermodynamics show a monotonic decrease in SFC benefit as overall pressure ratio 

increases [Ref. 104]. As such it was determined that for the size of engine under investigation in this 

study, CVC technology was not applicable, and could not be utilized to its full potential. This, combined 

with a low TRL, currently estimated to be 2, resulted in CVC/wave rotor technology not being considered 

for further investigation. 

B.2 Turboelectric Distributed Fans 

This type of engine architecture uses the core of a gas turbine engine to supply power to a generator, 

which in turn supplies power to a series of motors. These motors then drive propulsors, which may be 

distributed in any functional configuration around the airframe. This distributed propulsion system would 

ideally be combined with embedded, boundary-layer ingesting inlets to minimize the acoustic signature of 

the aircraft. Distributed propulsion can be used to create a more uniform velocity profile behind the 

vehicle, reducing drag and increasing propulsive efficiency [Ref. 106]. Theoretically, the gas turbine core 

can be driven at its optimum cycle point to minimize fuel consumption while supplying power to the 

generator. Electric drive trains offer improved energy efficiency over mechanical drive trains, along with 

greater system flexibility and lower system stress. The electric drive train would convert the mechanical 

energy from the engine to electrical energy through the integration of a high-speed generator, which 

would supply electricity to the distributed fans and other vehicle systems through the use of 

superconductors and power converters. According to a NASA presentation by Brown, current power 

converters/inverters operate at 95% efficiency, and are targeted to reach 99.8% efficiency [Ref. 107]. 

However, replacing the direct shaft connection between the gas turbine core and the propulsor with a 

generator and motor will always lead to a heavier propulsion system. This is especially true in light of the 

power densities of motors and generators required to make this system practical. In the coming decades, 

superconducting motors and generators can be utilized with larger power densities; however, such 

components require cooling systems, which in turn add additional weight to the system. 

Currently estimated at a TRL of 2, this technology has been deemed to be unrealistic in the N+3 

timeframe due to the complexity of the propulsion architecture, and the requirement for the more 

advanced components that comprise it. 

B.3 Turboshaft-Powered Direct-Drive Distributed Fans 

Turboshaft-powered direct-drive distributed fans use a single or multiple engines to provide power for 

up to four propulsors per engine. The enables ultrahigh-bypass engine performance, and allows core 

components to operate more ideally. This would be used with a partially- or fully-integrated propulsion 

NASA/CR—2010-216798 114
system. Multiple fans increase system redundancy, improving safety. This concept has also been 

suggested to be integrated for boundary layer ingestion, noise reduction, and high lift systems. As 

mentioned previously, distributed propulsion can be used to create a more uniform velocity profile behind 

the vehicle, reducing drag and increasing propulsive efficiency [Ref. 108]. 

Turboshaft-powered direct-drive distributed fans use a mechanical drive train instead of electric. The 

mechanical drive train employs several driveshafts and gearboxes to transfer energy from the engine to 

the distributed fans. A mechanical drive train in a dual-fan gear system shows 94% efficiency between the 

power of the fan and the power of the engine [Ref. 107]. Disadvantages of turboshaft-powered directdrive distributed fans include a weight penalty over conventional propulsion configurations. An electrical 

driveshaft configuration would have increased weight over its mechanical counterpart due to the need for 

generators, motors, inverters, coolers, and the rest of the superconducting system [Ref. 109]. Both 

systems would add increased complexity to airframe configurations, propulsion, and integration. 

Currently estimated to be TRL 2, this technology was not considered for further investigation due to 

its being out of scope of the N+3 timeframe. 

B.4 Nonturbine Distributed Fans 

Nonturbine distributed fans use distributed propulsion to achieve efficiency and drag benefits 

discussed for Turboshaft-Powered Direct-Drive Distributed Fans. The previous distributed propulsion 

technologies have a central engine to supply power to an array of fans, whereas the nonturbine distributed 

fan configuration uses separate independent power systems for each distributed fan. This allows for 

flexibility in the propulsive system, and allows for the use of alternative power systems or cycles. Diesel 

or biodiesel fuels could be used to reduce NOx emissions. Nonturbine distributed fans enable 

configuration flexibility through the use of shielding and placement. However, this propulsive 

configuration would add a significant amount of complexity to both the vehicle and propulsion systems, 

as well as require the propulsive system to be partially or fully integrated with the airframe. 

Currently estimated to be TRL 2, this technology was not considered for further investigation due to 

being out of scope of the N+3 timeframe. 

B.5 Hybrid All-Electric Propulsion 

Compared to other forms of hybrid propulsion, the all-electric version replaces the power source from 

the turbojet core to fuel cells, chemical batteries, solar cells, or any combination of electric-power 

generating options. Because these types of propulsion systems depend on propeller fans to generate thrust, 

noise is reduced dramatically. Depending on the power source, the remaining effects on the N+3 metrics 

varies, as do the system penalties. Driving the system with chemical batteries, for instance, simplifies the 

system tremendously; however, the weight of chemical batteries becomes prohibitive. The introduction of 

a fuel cell bank as the power source drives down LTO NOx emissions, but adds the complexity of 

additional components. For example, in order to supply the fuel cell with oxygen, a compressor and motor 

are needed. Furthermore, depending on the type of fuel cell, it may be necessary to include a jet fuel 

reformer or a supply of liquid or gaseous hydrogen. Of course, gaseous hydrogen storage is dangerous, 

and requires a large storage tank, and liquid hydrogen brings along with it the requirements of a 

refrigeration system. All forms of electric propulsion will also necessitate a power management and 

distribution system, with batteries to supplement the system and act as backup power supply. The benefit 

of all-electric propulsion is realized in the fuel reduction capability of the technology; however, the power 

requirements for takeoff drive the sizes of the components to infeasible proportions. 

For this reason, combined with the currently estimated TRL of 2, this type of propulsion system has 

been set aside as an impractical technology that will not be mature enough for aircraft applications in the 

N+3 timeframe. 

NASA/CR—2010-216798 115
B.6 Inlet Flow Control (Microvortex Generators, Jets, etc.) 

Several inlet optimization techniques exist for improving inlet flow field and conditions for inlets that 

are in, or near, the boundary layer. Recirculating fan bleed systems divert a portion of fan exit flow back 

to the inlet to increase the momentum of the boundary layer. This creates a more favorable boundary layer 

profile and a more uniform flow field, reducing distortion and preventing flow separation. Similarly, 

suction systems involve bleeding air from the inlet through the use of a porous surface in the region of a 

momentum-deficient boundary layer to prevent flow separation in embedded, blunt, and poorlyconditioned inlets. If the configuration allows, the inlet may be designed such that no inlet flow control is 

necessary. However, if the configuration is complex and volume is limited, a highly conformal inlet that 

is susceptible to flow separation or high levels of distortion may be necessary. 

Vortex generators require no engine bleed, and can be effective in reducing face distortion and 

preventing flow separation. Vortex generators create turbulent flow in the boundary layer, aiding in the 

formation of a more favorable boundary layer and increasing pressure recovery. In a study by Jirasek, 

vortex generators were found to have a maximum of 1.95% increase in pressure recovery. Vortex 

generators would result in a very small weight penalty [Ref. 110]. 

The pressure recovery associated with blowing and suction systems has the potential to be 

significantly higher than that of vortex generators, depending on the extent of integration and the level of 

blowing/suction. Engine performance gains bought through increased pressure recovery must be balanced 

against the penalties associated with engine bleed. Both blowing and suction flow control systems would 

result in a weight penalty, as complex ducting would be required, as opposed to vortex generators, which 

would result in a negligible weight penalty and require no engine bleed. None of these technologies 

would be required on a subsonic podded engine configuration, as extremely high inlet pressure recoveries 

are already achievable. 

The TRL varies per individual flow control technology. Due to the weight penalty associated with 

ducting, and the propulsive efficiency penalties due to engine bleeding, the flow control methodologies of 

recirculating fan bleed, blowing and suction, and inlet vortex generators were not investigated further. 

Inlet flow control was not considered for further investigation due to the vehicle operating in sub-sonic 

conditions which eliminates the formation of shocks in the inlet and therefore the need for a robust flow 

control system, such as blowing or suction. 

B.7 Wing Morphing (Piezoelectrics, Shape Memory Alloys, etc.) 

Wing morphing entails the ability to alter the shape of the wing, including aspect ratio, sweep, or 

airfoil design. Morphing allows for achievement of the optimal lift-to-drag ratio for each segment of the 

flight envelope by morphing the wing into the most optimized configuration. Wing morphing can be 

achieved through several systems, including pneumatics, shape memory alloy-based actuators, and 

piezoelectrics. These systems would be distributed throughout the wing surface, or in the sub-structure of 

the wing. Such systems would increase wing weight and complexity to the wing cavity, as well as remove 

wing volume that could be used for fuel storage. 

Wing morphing can be achieved by autonomously varying the camber of the wing through the use of 

flaps and symmetric aileron deflection, while flying away from the design condition. Because commercial 

aircraft spend the majority of flight time in cruise, there is little benefit to optimizing the wing geometry 

for climb and descent when considering the weight penalties involved. This technology would be more 

applicable to vehicles with varying flight conditions such as military aircraft that might have dash, loiter, 

and air-to-air engagement as part of their mission. 

The TRL varies per wing morphing technology. Due to these reasons discussed above, wing 

morphing was not further investigated. 

NASA/CR—2010-216798 116
B.8 All Electric (No Accessory Gear Box) 

Accessory gear boxes (AGBs) are currently used to convert mechanical energy from the engine into 

electrical energy from generators to be used by various subsystems of the vehicle, such as an AC 

alternator, DC starter, generator, and various pumps. Integration of electrical generators into the engine 

would eliminate the requirement for an AGB. Current AGBs are attached near the bottom of the engine, 

driving nacelle geometry to accommodate this added volume, and causing the nacelle to be wider and 

asymmetric. Removal of this component would lead to a more symmetric design, and reduced nacelle 

size, reducing drag and weight. If takeoff gross weight, fuel load, and wave drag are held constant, 

removal of the AGB could lead to a 0.6% increase in range and a 0.37% reduction in gross weight 

[Ref. 111]. Holding range constant, this would result in reduced fuel burn. If the vehicle were resized 

after the AGB removal, weight and fuel burn reductions would only increase. 

Currently estimated to be TRL 3, this technology was not considered for further investigation due to 

marginal improvements and incompatibility with the lightweight fan structure and cowl design. 

B.9 Structural Materials Laser Surface Treatment 

Advanced materials laser surface treatment uses a process known as laser peening, or laser shock 

peening. This process hardens metallic surfaces through the use of a powerful laser. Laser peening creates 

a highly compressible surface residual stress, which greatly improves structural properties, fatigue 

resistance, and life. This technology can be used to prevent stress corrosion cracking, overall corrosion, 

and pitting. Laser peening has demonstrated a smooth surface finish and good process control through 

testing and production. A laser is fired at a material that is covered with a dark coating, which absorbs the 

laser energy. The energy causes the coating to explode and produces a shock wave that propagates into 

the material, creating an array of compression indentations. This process is repeated until the entire 

desired area is treated. Laser peening has been, and is currently, demonstrated in industry with no known 

in-service problems. This technology is in full commercial production on two continents, including 

utilization in the F-22, P-3, and KC-10, as well as the Boeing 777, A340-500, and -600. The new Boeing 

787 uses laser peening on its Rolls-Royce Trent engine line, and has shown significant lifetime 

enhancement against failure and stress corrosion cracking. It was found, however, that laser surface 

treatment is not used for structural weight savings, but only to improve fatigue resistance and material life 

[Ref. 112]. 

Currently estimated to be TRL 5, this technology was not considered for further investigation due to 

having no significant vehicle weight, noise, or emissions benefits. 

B.10 Lightweight Foamed Composites 

Lightweight foamed composites are comprised of a sandwich structure with foam filler, based off 

nickel crystal polymer foam. Lightweight foamed composite has promising ballistic qualities, as well as 

being easy to machine. The advantages of foam include increased corrosion resistance and insusceptibility 

to water damage, giving it advantage over honeycomb core. Isotropic properties of foam enable increased 

durability. Lightweight foamed composites are seen as potential replacements for secondary structures 

that use honeycomb sandwich structures, including luggage bins, galley lower-wall panels, and ceiling 

panels. Based on internal analysis, foam core composites offer a lower strength-to-weight ratio compared 

to metallic and nonmetallic honeycomb core structures. However, as stated previously, foam core 

composites offer dimensional stability and increased durability, and may potentially leading to weight 

savings in select secondary structures. 

Currently estimated to be TRL 4, this technology was not considered for further investigation due to 

its small projected weight benefits. 

NASA/CR—2010-216798 117
B.11 Unsteady Circulation Control 

Unsteady circulation control is a subset of circulation control wing technology using pulsed jets of 

air, requiring less engine bleed, and reducing drag compared to steady circulation methods. This 

technology can enable short takeoff and landing distances, thus allowing for the down-sizing of lifting 

surfaces. Pulsed pneumatics are used to cyclically blow a sheet of air over a rounded trailing edge, taking 

advantage of the Coanda effect to delay separation of flow on the airfoil with extended flaps. Pulsed 

blowing effectiveness is largely dependent on the efficiency of the actuator system, as well as the 

response of the internal volume of gas prior to the jet exit. Ideal pulsed blowing requires a square 

waveform in wave response. However, as the complexity of the actuator system increases, this ideal 

waveform becomes more difficult to achieve. In addition, as the frequency of pulsing increases, blowing 

distortion increases, limiting the mass flow through the jet exit. At a duty cycle of 20%, unsteady 

circulation yields a lift coefficient increment of approximately 1.0. An equal lift coefficient with steady 

circulation requires approximately 92% more mass flow. At a fixed mass flow rate of 25 SCFM, the 

unsteady circulation control configuration yielded a 35% increased lift coefficient compared to the steady 

circulation control configuration [Ref. 113]. Weight penalties or benefits are configuration-dependent. 

Currently estimated to be TRL 2, this technology was not considered for further investigation due to 

its low TRL, combined with the result that traditional high-lift systems proved sufficient for this study. 

B.12 Water Injection 

This technology entails injecting high-pressure water into the combustor in order to decrease flame 

temperatures. Associated NOx emissions decrease due to the drop in combustion temperature. However, 

in the limit of excessive water use, less fuel will burn, and soot may form, affecting the performance of 

the combustor. This has been shown in computational studies performed by Fan et al [Ref. 114]. Water 

injection needs to be incorporated into a combustor in the design phase; retrofitting would be extremely 

difficult. In the implementation of this technology, the propulsion system requires that water be 

transported on the aircraft, as well as the associated injection mechanisms and control systems. 

Although it is currently evaluated at a TRL of 6, it has been deemed unnecessary to achieving the 

NOx goals of the N+3 study, in light of the success of the lean combustor technology outlined in Section 

5.3.3.18. 

B.13 Variable Pitch Fan Blades 

Controlling the pitch of fan blades would allow for the optimization of the fan pressure ratio at 

various operating conditions. This capability could remove the need for a thrust reverser and reduce 

windmilling drag [Ref. 115]. The incorporation of variable pitch fan blades would depend heavily on the 

methodology employed to generate the variable pitch. Any system that can control the geometry will add 

weight to the overall engine system. While experimental studies have shown that variable pitch fans can 

exhibit higher efficiencies than fans with variable inlet guide vanes, the mechanism to control the fan 

blade pitch would need to be compact enough to fit inside the fan hub [Ref. 116]. 

Currently estimated at TRL 5, this technology has potential to mature in the N+3 timeframe, but the 

necessity for the efficiency gains is outweighed by the complexity and weight of the system. 

B.14 Fuel Cell Auxiliary Power Unit 

A fuel cell auxiliary power unit (APU) would be used as a replacement for the current turbinepowered APU. When a vehicle is on the ground, and engines are off, vehicle power is supplied by the 

APU. Vehicles that spend a significant amount of time on the ground would see a significant fuel burn 

reduction from improvement in APU efficiency [Ref. 117]. The fuel cell APU uses hydrogen fuel in an 

electrochemical process to generate electricity. Due to integration constraints, hydrogen tanks are not 

NASA/CR—2010-216798 118
easily integrated or desirable, leading to the requirement of a jet fuel reformer. Historically, fuel cells 

have been extremely heavy, large in volume, and have demonstrated low power density. These properties 

have seen significant improvements in the past 30 years, and continue to improve. 

The use of a fuel cell APU in flight would improve engine efficiency and reduce fuel consumption by 

unloading a portion of the electrical load on the engines [Ref. 118]. Current aircraft are approximately 40-

45% efficient in converting fuel to electrical energy in the engine through a turbine, while fuel cell APUs 

can be up to 75% efficient in converting fuel to electricity during cruise. This increase in efficiency, and 

the relieving of portions of the electrical load from the engines, has been estimated to burn 40% less fuel 

in cruise for the purpose of producing electricity. However, it must be noted that the fraction of fuel used 

to produce electricity during cruise is minimal with respect to total aircraft fuel burn. 

Currently estimated to be at a TRL of 4, the fuel cell APU was not investigated further due to its low 

TRL, and the relatively low impact the technology would have on fuel reduction for the given mission. 

B.15 TO/C-Assist Propulsion Unit 

Takeoff and climb assist propulsion units are systems designed to enable STOL capabilities. 

Detachable autonomous propulsion units attach to the vehicle, and assist the aircraft during takeoff and 

climb. Upon reaching cruise altitude, the propulsion units detach and return autonomously to the airport. 

High-lift systems typically add weight and increase cruise drag while greatly improving field 

performance. This technology would enable the high lift benefits without the cruise drag or weight 

penalties. However, incorporating detachable autonomous propulsion units would add complexity to 

vehicle design, and increase the complexity of air traffic control, as propulsion units would need to be 

tracked and directed appropriately. In addition, they would increase the time and complexity of ground 

operations for maintenance, retrieval, and attachment of propulsion units prior to takeoff. Propulsion units 

that are expendable, jet-fueled assisted takeoff (JATO), could be considered, as such systems exist on 

many missiles, rockets, and targets. However, retrieval of units must be taken into consideration, as well 

as their impact on the surrounding air and ground space. 

Currently estimated to be TRL 1, this technology was not considered for further investigation due to 

traditional high-lift systems proving sufficient in this study, it’s extremely low TRL, and the added 

complexity to air traffic control operations, ground operations, and vehicle configuration. 

B.16 Microvortex Generators and Riblets 

Microvortex generators and riblets use passive means to affect the basic aerodynamics of the vehicle. 

Advances in CFD, nano-manufacturing, and greater understanding of boundary layer physics may soon 

enable the complete control of the aircraft flow field. Passive techniques such as riblets and microvortex 

generators do not require on-board support for power and pneumatics. Combining two or more passive 

drag reduction technologies could provide a multiplying effect. Riblet technology has demonstrated drag 

reduction, and tailored porosity, combined with riblets, is forecasted to enhance drag reduction. Riblets 

are designed for a single point design and can theoretically reduce friction drag by 8% at that point 

design. At off-design points, the benefits decrease, and at some flight conditions, riblets may cause a drag 

penalty. A small weight penalty would be present, as well as increased maintenance on surfaces where the 

technology is applied. 

Riblets have been integrated and demonstrated on commercial vehicles, including the A340, where 

they demonstrated a 1-1.5% decrease in fuel burn [Ref. 119]. A business jet equipped with riblets 

demonstrated a 6% reduction in parasite drag [Ref. 120]. In this report, the effects of riblet height and 

width were visualized. Additional correlations between shape and dimensions can be seen in Walsh 

[Ref. 121]. 

Microvortex generators have the capability to minimize uncontrollable airflow that causes boundary 

layer separation over the wings and flaps. Microvortex generators can be used to both increase lift and 

NASA/CR—2010-216798 119
decrease drag, enhancing vehicle performance during approach and landing. If the technology is used to 

reduce drag, less engine power is needed to produce equivalent lift on the aircraft, allowing for decreased 

landing and takeoff noise. Since microvortex generators can be easily stored in the flaps at cruise speed, 

there is no increase in cruise drag. Microvortex generators would have a small weight penalty associated 

with the technology. According to studies by NASA Langley Research Center, microvortex generators 

can result in a 10% increase in lift, 50% decrease in drag, and 100% increase in lift-to-drag ratio on a 

single flap configuration [Ref. 122]. 

The TRL varies per individual technology. Riblets have been flight tested and successfully 

demonstrated on commercial aircraft. For a short range vehicle, benefits from friction drag reduction 

would be less significant than a longer range mission during cruise. As mentioned previously, riblets at 

off-design points potentially experience an increase in drag. 

Both technologies are currently estimated at a TRL 3. Due to the relatively low percentage of fuel 

burn reduction and conventional high-lift systems proving sufficient, riblets and passive drag reduction 

were not investigated further. 

B.17 Conformal Polymer Solar Cell Power Augmentation 

Quantum dot nanoparticles embedded in polymer can be used to capture radiation in the extended 

electromagnetic spectrum beyond visible to infrared. Conformal polymers would coat the outside of an 

aircraft, and the captured electromagnetic energy would be converted to electricity to supplement or 

replace that generated by the gas turbine for use in auxiliary systems. According to Nadarajah et al., 

thicknesses on the order of 30 nm can be achieved with the polymer [Ref. 123]. However, energy 

conversion efficiencies are extremely low, on the order of 1%. While the weight penalty may be minimal, 

the electrical conversion rate is likewise minimal. Placing this polymer across 930 ft2

 of an aircraft would 

generate less than 1.6 hp. 

Although the technology is currently rated at a TRL of 3, it is believed that significant advances in 

energy conversion efficiency are required before it becomes practical to use as an alternate power source 

for commercial aircraft. 

B.18 Moveable Winglets 

Fixed winglets are common in modern commercial aircraft, as they reduce wing tip vortices and 

increase wing tip sectional lift. Both of these contribute to an increased lift-to-drag ratio and reduced fuel 

burn. However, fixed winglets are designed for cruise. Due to their fixed nature, they do not operate as 

well at off-design conditions. Moveable winglets could use shape memory alloy to reduce wing tip 

vortices and reduce fuel burn for all flight conditions, and produce lower bending loads than fixed 

winglets [Ref. 124]. In an interview with Wired, Airbus lead project engineer Askin Isikveren stated that 

fixed winglets optimized for cruise conditions vary from 3-5% reduction in fuel reduction. If moveable 

winglets were used, a 5% reduction in fuel could be maintained [Ref. 125]. Reduction in wing tip vortices 

and wing bending would further lead to a reduction in wing tip noise during takeoff and landing. 

Currently estimated at TRL 2, this technology was not considered for further analysis due to its long 

road to maturity, combined with the relative insignificance of wing tip noise compared to other airframe 

sources. 

B.19 Combined/Compound Cycles 

The integration of multiple thermodynamic cycles into a single engine core can drive up thermal 

efficiencies, with the potential for substantial increases in overall cycle efficiency. For example, cycles 

such as Otto, Brayton, Humphrey, Rankine, or Sterling cycles could be combined together to drive 

thermal efficiency gains. The incorporation of combined/compound cycles into aircraft engines could 

NASA/CR—2010-216798 120
eventually be performed on any scale for any type of aircraft, but would require a completely new engine 

architecture that would be a function of the cycles involved. Substantial complexity would be involved in 

such an engine, and it is estimated that the development of these new designs would take far greater time 

than the N+3 timeframe allows. 

Currently estimated at TRL 2, extensive investment is required before this technology can progress to 

the point of commercial application. 

B.20 Curved, Scarfed Inlet Ducts 

Scarfed inlet technology relies on an asymmetric inlet lip, protruding farther in the keel than in the 

crown of the inlet, to provide additional shielding in the direction of observers on the ground. The 

resultant flow field asymmetry has been shown to lead to extensive inlet distortion, which could 

theoretically lead to a negation of the acoustic benefits. However, modern flow control technology is 

more than sufficient to accommodate these levels of inlet flow distortion. 

This technology cannot be applied to embedded or boundary layer ingesting inlets, though these 

integration methods are generally sufficient in themselves to contain noise in the direction of the observer. 

In addition, scarfed inlets are known to exhibit reduced foreign object ingestion. Scarfed inlets will most 

likely be lighter to some degree than standard inlets, but may be more sensitive to angle of attack due to 

their asymmetry. 

Conservative weight estimates showing inlet weight reductions in the range of 5% are based on 

previous computational results by Abbott for scarfed inlets [Ref. 126]. Abbott has also shown that 

pressure recovery associated with a scarf inlet is in the range of 0.965 to 0.98 [Ref. 127]. Many acoustic 

studies have also sought to assess the noise benefits of scarfed inlet ducts [Ref. 128,129,130]. These 

studies have shown that this SPL reduction is approximately 5 dB in the direction of the observer. 

Currently estimated to be TRL 8, this technology was not considered for further investigation due to 

the weight of required flow control devices, and incorporation of the lightweight fan structure and cowl, 

discussed previously. 

B.21 Steady Fluidic Separation Control 

Steady fluidic separation control uses blowing or suction on the wing leading edge to energize lowmomentum flow in the boundary layer to delay or prevent flow separation [Ref. 131]. This enables the 

wing to operate at increased angles of attack, potentially shortening takeoff and landing field lengths. 

Steady fluidic separation control can be used as a replacement for leading-edge slats, which would reduce 

wing weight and noise. However, a blowing or suction system would require complex ducting, increasing 

weight, requiring engine bleed, and thereby decreasing engine performance. 

Currently estimated to be TRL 4, this technology was not considered for further investigation due to 

the estimated engine performance penalty, and due to traditional high-lift systems proving sufficient. 

B.22 Synthetic Jet Separation Control 

Synthetic jet separation control uses a small piezoelectric pump to convert acoustic oscillations into 

fluid motion. The piezo-electric pump is mounted inside a cavity near an aerodynamic surface. This 

cavity converges near the surface to focus acoustic oscillations. The converging portion of the cavity is 

oriented downstream, tangent to the flow and surface. This enables the pump to generate suction of the 

fluid upstream of the pump during the in-stroke, and tangential blowing of the flow aft of the pump 

during the out-stroke. 

Integration of a synthetic jet separation control system would add weight to the vehicle, in addition to 

requiring power to operate. While no engine bleed is required with the use of synthetic jet separation 

NASA/CR—2010-216798 121
control, power to operate the system must be generated by the engine. In a low Reynolds number 

experiment, McCormick found that the implementation of synthetic jet separation control could increase 

the stall angle approximately 37% and increase the maximum lift coefficient by 20% at a Reynolds 

number of 500,000 [Ref. 132]. In addition, the experiment demonstrated to be energy efficient when 

comparing system benefit to input power. 

Currently estimated to be TRL 3, this technology was not considered for further investigation due to 

the weight and system power penalty. 

B.23 Nanotube Liner 

A nanotube liner is an inlet liner designed to reduce tonal fan noise, similar to single and multi-DOF 

liners. Liners use small acoustic cavities to cause resonance of acoustic waves and dissipate acoustic 

energy. However, current liners are limited in the number of frequencies they can target. But nanotube 

liners could potentially provide a sound reduction system independent of frequency. Through the 

utilization of a honeycomb matrix that traps acoustic waves, energy is dispersed through viscous shear 

instead of resonance. The friction between the compressed gas of the acoustic wave and the honeycomb 

material slows the wave. Super alloys are the leading technology for larger, broadband noise dissipation 

applications, and are estimated to reduce aircraft engine noise by up to 30% [Ref. 133]. A tradeoff exists 

between the absorption capacity of the liner and the drag that is generated by the face sheet. Additionally, 

the weight and volume penalties associated with acoustic liners increases with thickness and area. 

Currently estimated to be TRL 2, this technology was not considered for further investigation due the 

technology being out of scope of the N+3 timeframe. 

B.24 Plane Jet Combustion 

The idea behind plane jet combustion is to inject fuel flow into a thin film distributed in space to form 

a plane of jet fuel. The aim of this method of combustion is to enhance combustion efficiency and lower 

NOx emissions. These low emissions derive from better fuel atomization, stemming from the increased 

surface area against which shear forces can act relative to standard injection techniques, combined with 

lower residence time. This type of combustion is also known to form strong recirculation zones on either 

side of the plane jet, providing a strong anchor for flame stabilization [Ref. 134]. 

While the current TRL of this technology is 5, it was set aside due to the successes achieved in 

lowering NOx emissions through staged combustion. 

B.25 Electric Subsystems 

Electric subsystems use an electrical converter, as well as other electrical components, as the main 

power supply for the aircraft. This power supply is able to power advanced radar, avionics, and a 

completely electromechanical flight control system. Aircraft dating from the F-14 to the B-2 have used 

power supplies connected to engines to produce a main power supply of 115 VDC, and have incorporated 

a hydraulic flight control system. The F-22 was the first aircraft to use 270 VDC as the main power 

supply, while still using a hydraulic flight control system. The F-35 and B-787 were the first two aircraft 

to fully integrate a 270 VDC power supply and an electromechanical flight control system. The 270 VDC 

power supply was originally designed to power large radar devices, but was adapted for the commercial 

market to power flight control system and avionics. 

An electromechanical flight system would increase system weight, but increase electrical efficiency, 

leading to improved engine performance and a reduction of fuel burned. Based on conservative estimates 

by Phillips and Wood, a vehicle with partially-electric subsystems would realize a subsystem weight 

reduction of 634 pounds compared to a conventional vehicle [Ref. 135]. However, a vehicle with fullyelectrical subsystems would see an increased subsystem weight. All vehicles in this study were sized for a 

NASA/CR—2010-216798 122
62,000 pound payload and a 2,600 nautical mile range. The vehicles with all-electric and partially electric 

subsystems showed an improvement in SFC of 1.41% of 0.895%, respectively. The vehicles’ fully and 

partially electric subsystems showed 1,421 and 3,113 lb reductions in ramp weight, and 569 and 2,141 lb 

reductions in operating empty weight, respectively. Due to the date of this report, and conservative 

estimates, benefits from fully- and partially-electric subsystems could be much greater in the N+3 

timeframe. 

Currently estimated to be TRL 8, this technology was not considered for further investigation. 

Improvements in weight and performance were not predicted to be significant. 

B.26 Third Engine Stream 

The third stream concept involves the introduction of an additional bypass stream into a standard 

turbofan engine, whose flow rate can be modulated. This can be achieved in a multitude of ways, 

including placing the inner fan duct downstream of two separate fans and the outer fan duct behind the 

upstream fan, or by having the outer fan duct supplied solely by the flow passing through the fan tip. 

Methods of modulating the flow, or varying the geometry of the outer fan duct, can be achieved through 

variable pitch fan blades, fan exit guide vanes, etc. This can be used to access higher bypass ratios than 

standard turbofans, with corresponding reductions in jet noise by shielding the high-speed, hightemperature core flow with slower outer fan duct flow. An additional flow path with variable geometry 

will allow for higher thrust levels at takeoff and climb and will allow for the maintenance of higher cruise 

efficiency. This separate flow path can also be used as a supply of air for flow control applications or 

thermal protection purposes. However, this can also result in larger engine diameters and additional 

complexity through the control system. 

While the currently estimated TRL of this engine architecture is 4, it was not chosen to be pursued for 

this particular mission due to the added complexity of supplying an additional fan duct with air, and the 

predicted system performance penalties associated with a larger engine. 

B.27 Alternative Hydrocarbon Biofuels 

Biofuels have the potential to replace fossil fuels, which is becoming important in light of the results 

of the scenario study. One of the major drivers behind biofuels is the reduced emissions over more 

conventional fuels. A study by the Baylor Institute for Air Science showed that the introduction of up to a 

30% fraction of cooking oil, plant-, and animal matter-based biofuel in Jet-A reduced NOx emissions 

[Ref. 136]. For fractions below 20%, no degradation in engine performance was recorded. Biofuels may 

be a cheaper alternative to kerosene-based aviation fuels in the long term, but currently will be more 

expensive. Many biofuels have already been tested in commercial engines; however, their characteristics 

must be considered relative to current aviation gasoline prior to implementing them into modern engine 

technology. Since emissions are shown in Section 6.2.3 to be extremely low due to the use of lean 

combustor technology, the need for alternative hydrocarbon biofuels has been set aside for the time being. 

Although the TRL is estimated to currently be around 4, the necessity for this technology is farther 

off in the future. 

B.28 Embedded/Conformal Inlets 

Embedded/conformal inlets reduce the size of the aerodynamic wake of the aircraft by ingesting a 

portion of the boundary layer and reducing the extent to which the inlet protrudes from the aircraft. This 

technology employs high area diffusion rates and large offsets which can enhance propulsion systems due 

to their light weight, tight and compact packaging, and low volume properties. However, due to the 

shortened inlet and the potential for severe duct turning, high distortion levels are common due to 

variation in total pressure in the incoming boundary layer. If the turning is extreme enough, boundary 

layer separation is possible in the inlet duct. 

NASA/CR—2010-216798 123
Embedded/conformal inlets would be utilized on embedded propulsion configurations. A decrease in 

inlet length and volume can potentially reduce the empty weight of a vehicle by 1.8%, and increase fuel 

capacity by 1.8%, from a normal shock inlet with serpentine diffuser [Ref. 137]. The inlet pressure 

recovery of this configuration would be approximately 9.6% lower than podded propulsion configuration. 

Inlet flow control such as vortex generators, bleed, or blowing could be used to reduce distortion and 

improve pressure recovery. However, adding such systems would add weight to the vehicle. 

This technology was not considered for further investigation due to the dismissal of configurations 

using embedded propulsion systems. 

B.29 Low-Noise Fan Blades 

NASA Glenn Research Center is developing a low-noise fan blade that has reduced weight with 

increased propulsive efficiency, named the quiet high-speed fan (QHSP). This blade is designed to reduce 

takeoff EPNL by 6 EPNdB [Ref. 138]. The fan blade is designed to channel free stream air into the core 

spinner, main core, and through the fan blade before exiting the fan blade trailing edge. This fan blade 

trailing edge flow energizes the airflow, and increases engine performance, while reducing fan noise. 

Noise levels produced by the fan are estimated to be proportional to the velocity of the fan tip speed to the 

fifth power. However, with the QHSP, NASA was able to design a high speed fan that reduces noise 

through forward swept fan blades and a leaned stator that creates a more favorable rotor-stator interaction, 

reducing this passing tone. It was found that the reduction in EPNL was largely due to the reduction in the 

passing tone and its respective harmonics. 

The current technology readiness level associated with low-noise fan blades is estimated to be TRL 2; 

however, due to the low TRL, combined with the added complexity of the blown fan blades, this 

technology was not investigated further. 

B.30 Fan Blade Flow Control 

As mentioned in Subsection B.29, high levels of tonal noise are caused by rotor and stator 

interactions at the blade passing frequency and its harmonics. The use of flow control in this region could 

result in reduced engine noise. Air is ducted through the fan blades and ejected through the fan blade 

trailing edge, thus suppressing unsteady fan blade and stator interactions by reducing viscous boundary 

layer losses. Blowing can be applied in the fan blades in several configurations: every-blade trailing edge 

blowing (TEB), alternating fan blade trailing edge blowing (ATEB 1x1) and applying trailing edge 

blowing to two adjacent fans, skipping the next two, and applying trailing edge blowing to the next two 

(ATEB 2x2). The latter two cases result in reductions of peak tones but introduce several new tones. 

However, this new acoustic spectrum is easier to control by acoustic liner treatment than regular TEB. 

Figure B-1 shows the reduction in sound power as a function of blowing rate, as a percentage of inlet flow 

[Ref. 139]. Using trailing edge blowing in the absence of stators, the far-field noise was found to be 94.8 

dB, placing a lower bound of the achievable sound power level by removing all rotor-stator interaction 

noise. TEB reduces the inlet and aft duct sound power levels by 7% and 6.9% with blowing rates of 1.5% 

and 1.52%, respectively. ATEB 1x1 reduces the inlet and aft duct sound power levels by 4.3% and 5% 

with blowing rates of 0.9% and 0.9%, respectively. Reduced amounts of blowing result in increased 

engine performance. Additionally, it was found that using TEB, ATEB 1x1, and ATEB 2x2 increases 

acoustic liner sound reduction by 31%, 50.7%, and 58%, respectively, showing that fan trailing edge 

blowing increases acoustic liner effectiveness. 

NASA/CR—2010-216798 124
Figure B-1. Sound Power vs. Blowing Rate (% Inlet Flow Rate) for 

(a) Inlet Duct and (b) Aft Duct [Ref. 139] 

According to Fite et al, the fan efficiency of an engine with 2% blowing is equivalent to the efficiency 

of a baseline fan with no trailing edge blowing [Ref. 140]. There would be a weight penalty and added 

complexity associated with the engine, as ducting would need to be installed. Currently estimated to be 

TRL 4, this technology was not considered for further investigation due the estimated degree of benefit 

being outweighed by the added weight and complexity. 

B.31 Conformal Load-Bearing Antenna Structures 

Conformal load-bearing antenna structures (CLAS) embed sensors such as antennas and avionics 

where traditional methods use bolt-on methods. Traditional avionic sensors are able to be replaced with 

active or passive sensors that allow for controllable and reconfigurable antennas. CLAS eliminates bolton sensors and structural cutouts associated with these methods, improving structural efficiency, and 

enabling the consolidation of single-function sensors into multifunction installations. This allows for 

potential weight reduction and volume reduction in the vehicle. CLAS offers potential drag reduction by 

allowing for a more continuous skin surface. Based on internal research and studies, CLAS could reduce 

vehicle empty weight between 200 and 1400 lbs, reduce acquisition cost between $0.5 million and $4.0 

million, and may offer potential noise and drag improvements through a more continuous skin surface. 

This technology has a varied TRL, based on configuration and extent of integration, but was not 

considered for further investigation due to the small magnitude of weight, drag, and noise reduction. 

Commercial aircraft contain significantly fewer sensors than military or surveillance vehicles. As a result, 

a commercial vehicle that uses CLAS would see benefits toward the lower end of the ranges discussed 

above. 

B.32 Upper Surface Exhaust Nozzle and Flap 

The upper surface exhaust nozzle and flap system uses pylon-mounted engines located on the wing. 

High-momentum exhaust from the engines is vectored downward which prevents flow separation during 

large flap deflection angles and provides super circulation [Ref. 141,142]. This increased flap deflection 

and super circulation would lead to an increase in lift coefficient during takeoff and landing to improve 

vehicle field length performance. Integrating the vectoring nozzle and ducting would increase wing 

weight and complexity, and decrease usable wing volume. Vehicle parasite drag would increase for all 

segments of mission, increasing fuel burn. Additional penalties associated with mechanical thrust 

vectoring can be found in the Subsection B.71. Figure B-2 illustrates a comparison between an upper 

surface exhaust nozzle and flap configuration and a conventional flap configuration. Upper surface 

exhaust nozzle with flaps shows significant ESTOL characteristics compared to the traditional flap 

configuration. 

(a) (b)

NASA/CR—2010-216798 125
Currently estimated to be TRL 3, this 

technology was not considered for further 

investigation. For the mission requirements, 

traditional high-lift proved sufficient. 

B.33 Lower Surface Exhaust 

Nozzle and Flap 

Lower surface exhaust nozzle and flap is 

similar to upper surface exhaust nozzle and flap, 

but instead of mounting the engine on the upper 

surface, the engine is mounted beneath the 

wing. Externally blown flaps (EBF) energize 

flap flow with the use of engine exhaust. EBF 

delays or prevents flow separation on high lift 

systems and allows for greater flap deflection. 

These effects contribute to an increase in lift, 

reducing field length and allowing for steeper 

climb and descent trajectories. This, in turn, 

limits the time and area exposed to aircraft 

noise. According to two-dimensional analysis 

by Griffin et al, at a 0° angle of attack, lower 

surface EBF produced a 157% and 273% 

increase in lift coefficient for flap deflections of 

20° and 30°, respectively, during takeoff 

[Ref. 143]. However, lift-to-drag ratio decreased 

71.3% and 68.7%, respectively. At a 0° angle of 

attack, lower surface EBF produced a 220.8% 

and 490.6% increase in lift coefficient for flap 

deflections of 30° and 35°, respectively, during 

landing. However, lower surface EBF reduced 

the lift-to-drag ratios 65.4% and 52.8%, 

respectively. On the other hand, three-dimensional results have shown a decrease in lift performance. 

Based on three-dimensional analysis, at a 0° angle of attack, lower surface EBF produced a 6.9% and 

37.8% increase in lift coefficient for flap deflections of 20° and 30°, respectively, during takeoff. Lift-todrag ratio decreased 36.2% and 46.1%, respectively, as a result of lower surface EBF. At a 0° angle of 

attack, lower surface EBF produced a 20.8% and 128.8% increase in lift coefficient for flap deflections of 

30° and 35°, respectively, during landing. However, lower surface EBF reduced the lift-to-drag ratios 

24% for 30° flap deflection and increased the lift-to-drag ratio 12% for 35° flap deflection. Benefits of 

EBF vary per configuration, and decrease as wing span increases. Based on internal research, EBF would 

increase propulsion system weight approximately 6%, and increase scrub drag by 0.0001. 

Figure B-2 illustrates a comparison between an lower surface exhaust nozzle and flap configuration 

and a conventional flap configuration. Lower surface exhaust nozzle with flaps shows significant ESTOL 

characteristics compared to the traditional flap configuration. 

Currently estimated to be TRL 3, this technology was not considered for further investigation, as 

traditional high-lift devices proved sufficient to meet the field length requirements. 

B.34 Fluidic Thrust Vectoring for Upper Surface Exhaust Nozzle and Flap 

Fluidic thrust vectoring for upper surface exhaust nozzle and flap is similar to the upper surface 

exhaust nozzle and flap described above. Fluidic thrust vectoring provides a smaller weight penalty and 

Figure B-2. Upper Surface Exhaust Nozzle and

Flap Lift Comparison 

NASA/CR—2010-216798 126
an improved thrust efficiency compared to mechanical thrust vectoring as described in the Subsection 

B.71. Currently estimated to be TRL 3, this technology was not considered for further investigation. For 

the mission requirements, traditional high-lift proved sufficient. 

B.35 Single DOF Liner 

Single degree-of-freedom (DOF) liners are a type of inlet acoustic liner generally comprised of 

acoustic absorbing material covered by a perforated face sheet. The absorbing sections of the liner may be 

divided up into smaller cells designed to target specific tonal frequencies. In the inlet liner design process, 

face-sheet resistance and cavity depth are optimized to minimize noise in the inlet. Noise in the inlet is 

generated primarily by fan tonal noise through the interaction between rotor blades, stator blades, and the 

rotor pressure field. At subsonic speeds, however, inlet noise is dominated by rotor-stator interactions. 

The most common type of inlet acoustic liner is a single degree-of-freedom liner, referring to the 

number of frequencies that can primarily be targeted. Acoustic impedance is dependent on liner 

resistance, cavity reactance, and face-sheet mass reactance [Ref. 144]. A tradeoff exists between the 

absorption capacity of the liner and the drag generated by the porous face sheet. Additionally, the weight 

and volume penalties associated with acoustic liners increases with the number of degrees-of-freedom. 

Based on analysis, it was found that a double DOF inlet liner provided increased noise reduction 

compared to a single DOF liner. As such, this technology, currently estimated to be TRL 9, was not 

considered for further investigation. 

B.36 Active Drag Reduction 

Active drag reduction focuses on reducing the vehicle boundary layer, leading to a reduction in skin 

friction drag. Transition control and laminar flow encounter problems for very large applications and 

regions with increased local Reynolds numbers, such as the fuselage, where as much as half of the skin 

friction on the vehicle occurs. For a successful active device, the energy needed to control and reduce the 

turbulent boundary layer must not exceed the energy saved from turbulent boundary layer reduction. 

While active flow control systems require energy to function, they offer potential larger reductions in drag 

over passive systems. 

Past CFD studies have shown drag reductions of 25-30% from suction and blowing [Ref. 145], 71% 

from tangential blowing [Ref. 146], 40% from spanwise wall oscillation [Ref. 147], 20% from utilization 

of counter-rotating streamwise vortices, and 50% from utilization of colliding spanwise jets [Ref. 148]. 

However, CFD studies tend to overestimate system performance by allowing for control at each grid 

point, when this is not possible in a realistic active control environment. In terms of experimental studies, 

a 45% reduction in drag has been experimentally measured through spanwise oscillation [Ref. 149] and a 

7% reduction in drag has been measured through the use of an array of spanwise synthetic jets [Ref. 150]. 

The current TRL varies per individual technology. These technologies were not considered for further 

investigation due to the relatively low TRLs, weight penalties, and operating energy restrictions 

associated with each technology. Several technologies would require large amounts of complex ducting 

which would further increase weight and possible engine bleed which would decrease engine 

performance. 

B.37 Multi-DOF Liner 

Multi-DOF liners are built upon the current state-of-the-art single DOF acoustic liners for reducing 

tonal fan noise. Multi-DOF liners contain multiple layers of absorbing material and additional perforated 

sheets separating it into other layers. Similar to single DOF liners, multi-DOF can have the absorbing 

sections divided up into smaller cells designed to target specific tonal frequencies. Multi-DOF liners 

allow for attenuation which targets multiple different tonal frequencies. A tradeoff exists between the 

absorption capacity of the liner and the drag that is generated by the porous face sheet. Additionally, the 

NASA/CR—2010-216798 127
weight and volume penalties associated with acoustic liners increases with its number of DOF. Based on 

analysis, it was found that a double DOF inlet liner proved superior in noise reduction and overall system 

performance compared to single- and multi-DOF liners. As such, this technology, although currently 

estimated at TRL 8, was not considered for further investigation. 

B.38 Oil-less Engine 

Oil-less engines use magnetic bearings instead of tradition oil-lubricated bearings. These magnetic 

bearings could potentially offset the weight reduction seen by the removal of conventional bearings, 

though a slight weight increase would likely be present at a local level [Ref. 151]. Magnetic bearings 

would eliminate friction inherent in gas turbine engine systems, improve shaft rotation, and reduce fuel 

consumption. Engine efficiency is also increased through larger clearances associated with magnetic 

bearings, and the elimination of engine operating temperature restrictions normally found with other 

forms of bearings and liquid lubricants. In addition to efficiency improvements, magnetic bearings would 

remove one of the main factors limiting single-mission duration and engine lifetime. Though magnetic 

bearings would slightly increase weight at a local level, their use would lead to the elimination of 

lubrication plumbing system, and simplification of the engine casing and support, leading to a potential 

10% decrease in overall engine weight [Ref. 152]. 

Currently estimated to be TRL 2, this technology was not considered for further investigation due to 

its need to mature. 

B.39 Low Tip Speed Counter-Rotating Ducted Fan 

This engine architecture concept entails using two separate fans with reduced tip speed compared to 

conventional turbofans. Similar to an open rotor concept, the fans will be counter-rotating, necessitating 

either a gearbox system or separate shafts and turbines to drive each fan. This will unavoidably add 

weight to the system in either configuration. Though the ducted architecture can allow access to higher 

flight Mach numbers through the diffusion of inlet air, and noise reduction compared to its unducted 

counterpart, it cannot achieve the fuel reduction potential of unducted propfans. Furthermore, while 

efficiency gains and noise reduction can be bought through the dual-level compression and reduction in 

fan tip speed, the required lower fan diameter comes with the added penalty of reducing bypass ratio, 

providing a further hindrance to fuel savings. In addition, ducted propfans are reported to have fan 

pressure ratios lower than more conventional turbofan engines, putting the engine at risk of surge. These 

conflicts with high efficiency necessitate the implementation of variable pitch fan blades, discussed 

elsewhere in this report. 

At a current TRL of 3, this technology was not considered for further pursuit due to the combined 

effects of reduced bypass ratio and reduced fuel consumption relative to the alternative configurations, 

specifically, the open rotor and three-shaft turbofan. 

B.40 Wheel Caps 

Wheel caps are designed to cover the hubs of landing gear wheels to reduce the separation that occurs 

in this cavity region while the landing gear is deployed. When present, this separation leads to lowfrequency airframe noise due to cavity flow. Assuming there is a way to affix the wheel caps to the 

wheels, this technology could be applied to any aircraft, and should generate positive results in terms of 

noise reduction. However, the added weight must be taken into consideration. 

As reported in Dobrzynski et al, wheel caps provide a benefit of 0.5 dB reduction in OASPL in the 

forward radiation direction, but are especially beneficial in the 500 Hz range [Ref. 66]. In addition, the 

work of Soderman et al shows a benefit of 0.6 dB due to the addition of a wheel cap [Ref. 153]. Based on 

a 3 foot landing gear tire, an estimated 10 pound wheel cap would be required for each tire. With two tires 

NASA/CR—2010-216798 128
on the nose gear and two on each of the main landing gears, this would add up to a total of 60 pounds of 

additional landing gear weight. 

Currently estimated to be TRL 6, this technology was not considered for further investigation due to 

the small magnitude and narrowband nature of the noise attenuation benefit versus the weight penalty that 

would be incurred. 

B.41 Wheel Gap Filler 

As described by Smith and Chow, an axial gap between two landing gear generates noise due to the 

impingement of the turbulent wake from the leading wheel on the leading edge of the rear wheel 

[Ref. 154]. The majority of this noise is directed towards the ground, as described by the computational 

results of Lazos [Ref. 155]. A wheel gap filler consists of a solid structure designed to occupy the volume 

between the two axially spaced wheels, in order to prevent the turbulent wake of the leading wheel from 

impinging on the trailing wheel. While this structure necessitates adding weight to the landing gear, it 

must be weighed against the noise benefit that can be achieved in terms of EPNL. Since wheel gaps are 

generally associated with a broad mid-frequency peak, these are the frequencies that would be affected by 

the inclusion of a wheel gap filler. 

Although the estimated TRL of this technology is 6, due to the relative contribution of the wheel gap 

source to the overall landing gear noise, this technology was not considered for further analysis on the 

configurations under investigation in this study. 

B.42 Acoustic Excitation 

Acoustic excitation is designed to promote mixing through the use of different tonal frequencies that 

induce turbulence in excited jets. It has been shown that significant modifications can be made on vortex 

formation on large scale structures through acoustic excitation. These induced vortices can lead to 

increased mixing, causing the noise source region to become more compact, smoothing out the jet 

velocity profile, and reducing the overall jet plume length, leading to reductions in far-field noise 

[Ref. 156]. Mixing of the jet flows is similar in effect to chevrons, without the structure required for 

chevrons, and the drag penalty associated with the technology. 

Currently estimated to be TRL 4, this technology was not considered for further investigation due to 

the small magnitude and narrowband nature of the noise attenuation benefit and low TRL. 

B.43 Brush-type Trailing Edges 

Brush-type trailing edges employ flexible fibers emanating from the trailing edge of a lifting surface. 

These bristles serve to alleviate the bluntness of the trailing edge, which would otherwise result in largescale vortex shedding. As a result, the noise associated with these shed vortices is mitigated to some 

degree. Furthermore, the flexibility of the fibers allows them to remain unloaded and adapt to different 

flow conditions. This technology can be applied to trailing edges of any lifting surface in order to mitigate 

vortex shedding. Its passive nature suggests that it will have a minimal weight penalty associated with it, 

and no added complexity in terms of electronics or control systems. Failure of this technology would 

consist of the detachment of the bristles on the trailing edge of the lifting surfaces. Depending on the 

attachment methodology, the entire brush-type trailing edge could be removed, or only certain bristles. In 

either case, the portion of the trailing edge that becomes detached would experience vortex shedding, 

leading to increased airframe noise. As reported in Herr and Dobrzynski, a properly designed trailing 

edge brush can show a noise reduction in excess of 10 dB (in a limited frequency range proportional to 

the flow velocity and trailing edge thickness), citing the ability of flexible brush fibers to adapt to the flow 

and be minimally affected by flow disturbances by remaining unloaded for different incidence conditions. 

This is in contrast to trailing edge serrations, which can only achieve noise reductions on the order of 2-

3 dB [Ref. 157]. Aerodynamic effects of including brush-type trailing edges were studied computationally 

NASA/CR—2010-216798 129
by Ortmann and Wild [Ref. 158]. While the noise reduction at the target frequencies can be significant, 

the broadband frequency reduction potential of the technology is severely limited. Additionally, the 

targeted acoustic source is among the most negligible contributors to the overall EPNL of the aircraft. 

Research has also shown a potential to reduce drag on the wing by controlling the trailing edge flow; 

however, this is accompanied by a corresponding decrease in its lift coefficient. 

Because only specific frequencies can theoretically be targeted with this type of trailing-edge device, 

even though its current TRL stands at 4, the ability of this technology to dramatically affect the noise 

footprint of the aircraft is minimal. For this reason the technology was set aside from further 

consideration for the duration of this study. 

B.44 Trailing Edge Serrations 

Similar to the trailing edge serrations researched for control of flap noise, serrations can also be 

applied to the trailing edges of flaps or the main wing element. Trailing edge serrations would involve 

cutting or constructing V-shaped grooves into the flap or wing material. In the presence of serrations, the 

periodic vortex shedding from the trailing edge of the wing element that would typically occur becomes 

less coherent in the spanwise direction. Instead it would be replaced by longitudinal vortices that form 

from the serrations. This can be seen in the oil-flow visualizations produced by Knepper and Garry to 

delay flow separation on the upper surface of the flap [Ref. 159]. Further wind tunnel measurements by 

this group showed that an optimally placed flap, with well-designed serrations on the wing trailing edge, 

can see an L/D benefit of up to 17% due to this effect. However, this benefit only corresponded to a 

change in wing lift coefficient of less than 0.1. Furthermore, the noise benefit is expected to target only 

the wake shedding frequency associated with the finite thickness trailing edge, replacing it instead with 

higher frequency noise from the serrations. 

Currently estimated at TRL 4, this technology has been set aside as its projected benefits are small 

compared to the overall acoustic field surrounding the aircraft. 

B.45 Chevrons 

Chevrons are triangular serrations integrated into the exhaust nozzle trailing edges of turbofan 

engines. These notches are designed to induce streamwise vorticity into the shear layers. This vorticity 

promotes the mixing of the jet streams, reduces the overall jet plume length, and leads to reductions in 

far-field noise [Ref. 160]. This technology is the current state-of-the-art in medium- to high-bypass ratio 

turbofan engine jet noise reduction, and imposes a negligible weight penalty to the engines. However, 

chevrons penetrate slightly into the exhaust flow, leading to thrust penalties at cruise. For this reason, 

some studies are looking into deployable chevrons by means of shape memory alloys, etc., which can be 

deployed for take-off noise reduction and retracted during cruise. As chevrons are designed into the 

structure of the engine, no failure modes are associated with them. Experiments have shown that 

approximately 3 dB reduction can be achieved with a chevron nozzle, mostly at lower frequencies, while 

incurring at most a 0.55% thrust penalty at cruise [Ref. 161]. Because the frequency reduction potential is 

limited to lower frequencies, and the associated thrust penalty is carried through the entire cruise portion 

of the mission in the absence of advanced control systems, this technology was excluded from further 

consideration. 

While the current TRL of this technology is at 9, and some aircraft have implemented them, the 

emphasis on fuel burn savings derived from the scenario study has led to this technology being set aside 

for the remainder of this study. 

B.46 Combined Steam Cycle 

A derivative of the combined/compound cycle technology described above, a combined steam cycle 

can described a variety of different engine architectures, though in general the features remain the same. 

NASA/CR—2010-216798 130
In one example, portions of the hot jet exhaust are utilized to support a closed Rankine cycle. The hot jet 

exhaust is passed through a heat exchanger and used to vaporize liquid water into superheated steam. In 

turn, this steam is passed through a turbine which is used to drive a shaft. This shaft can then be employed 

to create excess power through a generator or to drive a propeller to produce additional thrust. The twophase mixture leaving the steam turbine is then passed through a condenser, after which it becomes liquid 

water again. At this point it is forced through a pump and again passed though the heat exchanger. This 

type of system architecture introduces a variety of additional components to the engine: the water pump, 

heat exchanger, condenser, steam turbine, and either a generator or propulsor. Furthermore, there is a 

need for additional piping/tubing to transport the water between these components. In an industrial 

application where this has been implemented, steam tubes can be excessively large. Because of this their 

direct translation into an aircraft application can be prohibitive, unless smaller diameter tubing can be 

utilized. 

For aircraft engine applications, the estimated TRL of the combined steam cycle is only 2. More 

complete use of the heat generated by the gas turbine cycle increases the overall efficiency of the engine; 

however, the addition of a steam cycle into the system would require a complete redesign of the engine 

architecture. Furthermore, weight penalties become excessive in light of the large number of components 

required to perform the cycle. While theoretical cycle analyses have shown the potential of such a cycle to 

reduce specific fuel consumption by up to 20% in an uninstalled engine [Ref. 162], this technology has a 

long road to maturity before it can be considered further as a potential propulsion system for aircraft 

applications. 

B.47 Linear Acoustic Liner 

Similar to the construction of single degree-of-freedom acoustic liners described above, linear 

acoustic liners consist of a cellular material, such as honeycomb, sandwiched between a solid back sheet 

and a perforated front sheet. However, with linear liners this perforated front sheet generally consists of a 

bonded woven metal wire cloth. For the liner to be considered linear, the perforations on the front face 

sheet must be sufficiently small to force the acoustic impedance of the cavities to be insensitive to the 

magnitude of the flow rate across the liner. It is reported that this critical diameter is 10 μm [Ref. 163]. 

For this reason, it is far more practical to develop linear liners by using woven cloth. Experimental 

measurements, such as those reported by Drouin et al, explain that flow through the porous holes of a 

linear liner is completely dominated by viscous forces [Ref. 164]. As such, this type of liner will exhibit a 

skin friction coefficient that is largely insensitive to sound pressure level. Liners consisting of larger 

perforations, on the other hand, tend to experience increases in skin friction as the sound pressure level 

rises due to the effect of pressure fluctuations on the flow inside the cavities. For this reason, linear liners 

are more capable of operating successfully over a wide range of flight conditions. A downside to this 

design, however, is the impracticality of using woven metal cloth in the exhaust nozzle due to the high 

temperature flow. This limits the use of this type of liner to inlets and fan ducts. In addition, for larger 

engines it is limited in its sound pressure level attenuation compared to other liner designs for community 

noise flight conditions. 

Though the current TRL for this technology is 9, it was decided that a larger degree-of-freedom liner 

would be optimum for the configuration being explored in this study. 

B.48 Four-Shaft Turbine 

Like the three-shaft turbofan, a turbofan driven by four shafts has the added independence of 

rotational components with respect to one another. This allows for more of the turbomachinery to operate 

at ideal speeds, while stage matching increases thermal efficiency. The additional shaft can be utilized for 

additional fan stages, compressors, power turbines, etc. 

NASA/CR—2010-216798 131
Currently estimated at TRL 1, this technology is predicted to increase the weight of the engine for 

only a minor benefit in fuel consumption relative to a three-shaft turbofan. For this reason, combined with 

the long road to technological maturity, the technology was set aside from further consideration. 

B.49 Active Noise Cancellation 

Active noise cancellation is based on destructive interference of airframe noise by the introduction of 

an out-of-phase power source. Methods by which this can be achieved are to employ either a highpowered acoustic source in the form of a loudspeaker, or to generate an out-of-phase vibration into the 

noise-radiating elements of the aircraft. This methodology allows for the active control of noise 

generation in any portion of the aircraft into which it can be installed. Both methods require the 

measurement of generated noise to properly tune their destructive output. Active noise cancellation 

systems must be built into the elements of the aircraft that will be expected to generate noise. This will 

involve not only the actuation mechanism to force the vibration, but also the electrical and control 

systems. Power considerations must also be taken into consideration. Using an acoustic source to 

destructively interfere with the airframe noise will demand expensive power requirements, but may be 

easier to implement in existing aircraft. Weight and system complexity must be considered. Noise 

reduction from active noise cancellation varies per component and noise cancelation method. 

Based on previous research, active noise cancellation can achieve duct noise reductions of up to 20 

dB for frequencies up to 500 Hz, fuselage noise by more than 13 dB at certain frequencies, and fan noise 

reductions between 10-20 dB. There are several methods used in active noise control, including near 

source, which is desirable due to its ability to cancel noise at the source and prevents noise propagation. 

However, this is difficult to implement due to the location of many noise sources in extreme 

environments. Enclosures are another method that can be used to either prevent noise from propagating or 

preventing noise from entering a confined space. They allow for nonintrusive equipment placement, and 

allow the control of a well-defined space. However, enclosures can become large and complex. Singlepoint noise control aims to cancel noise at a single point. Applications are limited, but high performance 

results are easily achievable. Single-point uses a microphone to pick up a tone and uses a secondary 

source to emit an inverted signal. The last method for active noise control is free-field, which uses 

secondary sources to surround a primary source to prevent noise propagation or to create a region into 

which no noise can enter. Secondary sources must be placed no more than half a wavelength together, 

which is generally not possible. 

Currently estimated at TRL 2, this technology was not investigated further due to the added weight 

and complexity, decreased energy efficiency, and low TRL. 

B.50 Landing Gear Assembly Plasma Fairings 

Landing gear assembly plasma fairings consist of electrodes which introduce energy into the flow 

field surrounding the landing gear assembly. The “fairing” typically would take the form of a single 

dielectric barrier discharge (SDBD) plasma actuator and a power supply. The addition of this energy into 

the flow field delays or prevents separated flow in the bluff body wake of the landing gear, which is a 

primary source of landing gear noise. This technology could be applied to any unfaired landing gear 

system which meets the spatial requirements associated with the implementation of the plasma generating 

devices. There would also be power requirements involved with the use of such technology, and a slight 

weight penalty. Failure of the plasma system would simply reduce the landing gear to its nominal dressed 

state, in which there is extensive vortex shedding noise. In addition, the baseline drag of the system would 

be restored. There are no expected modes of failure that would cause a significant disruption of aircraft 

activity. Experimental work conducted at Notre Dame, as reported in Thomas et al, shows the application 

of landing gear plasma fairings for landing gear noise reduction [Ref. 165]. A subsequent work by the 

same group shows the source SPL reduction from a vortex shedding cylinder, and it was found that up to 

13.3 dB reduction in noise could be achieved by pulsing at the Strouhal number associated with the 

NASA/CR—2010-216798 132
vortex shedding [Ref. 166]. In addition, a drag coefficient reduction of 30% for the landing gear is 

assumed based on related work on physical landing gear fairings (e.g. Quayle et al [Ref. 65]). 

While the potential of this technology to address the challenges of the N+3 metrics exists for a 

minimal weight penalty, it was determined that at a current estimated TRL of 2, this technology could not 

be made viable in the timeframe of aircraft entering service in the range of 2030-2035. However, given 

another decade or so to mature, this technology could prove feasible for future airframes to use in meeting 

stringent noise regulations. 

B.51 Slat Cove Fillers 

Upon deployment of leading edge slats, a slat cove is generated on the lower side of the slat where it 

conforms to the wing in the undeployed state. This technology seeks to fill this gap with a solid body to 

prevent flow separation and result in a more streamlined flow between the leading edge of the main wing 

and the lower surface of the slat during takeoff. At the same time, it is desired to maintain the overall 

aerodynamic performance for the high-lift devices in which they are to be implemented. The main 

purpose of this device is to minimize vortex shedding resulting from lower side flow separation, which 

leads to broadband noise generation. 

This technology could be applied to any wing which employs slats, though in an end-product form, it 

would need to be deployable so that the slat and wing would still be able to conform to one another. There 

will be a slight weight penalty due to the additional material, and any deployment mechanism that must 

be added could add complexity to the system. In the experimental work by Kolb et al [Ref. 167], it is 

shown that a slat cove filler can achieve an average slat SPL reduction of around 4.6 dB across the 

frequency range. This noise reduction is greatest in the range 1-3 kHz. 

The current technology readiness level associated with slat cove fillers is estimated to be TRL 4; 

however, due to their noise reduction spanning only a slight portion of the frequency spectrum, combined 

with the relative importance of the slat noise when compared to other airframe noise sources, this 

technology was not considered for further analysis. 

B.52 Continuous Mold-Line Linkages 

Continuous mold-line linkages are designed to reduce noise generated from slat and flap side edge 

vortices. Noise is reduced through the blending of the wing and slat leading edges, as well as the wing 

and flap trailing edges. This blended surface greatly diffuses the vortices caused by slat and flap linkage 

gaps, and significantly reduces aircraft noise. According to Berton et al, continuous mold-line linkages 

are estimated to reduce flap noise by 8 dB [Ref. 168]. Noise reduction is constant across all one third 

octave band frequencies, ranging from 0.1 to 10 kHz and across all directions and airspeeds. Continuous 

mold-line linkages would increase slat and flap system weight a minimal amount. 

The current technology readiness level associated with continuous mold-line linkages is estimated to 

be TRL 3; however, due to the low TRL combined with the relative importance of flap noise compared to 

other airframe noise sources, this technology was not considered for further analysis. 

B.53 Flap Tip Fences 

Flap tip fences are flat plate obstructions employed at the inboard and/or outboard portions of wing 

flaps. They can be aerodynamically contoured depending on the platform on which they are used, 

extending below the lower surface of the flap. Their purpose is to prevent the shedding of streamwise 

vortices from the edges of the flap caused by the discontinuity in the wing surface upon deployment of the 

flap. These vortices have been associated with high levels of airframe noise during landing. This 

technology can be applied to any wing flap, and generally can be retrofitted to any wing flaps that do not 

currently implement them. However, there will be a slight weight penalty due to their implementation. In 

NASA/CR—2010-216798 133
Soderman et al [Ref. 153], source noise reduction using a flap tip fence was in the range of 1-3 dB. In 

Horne et al, a reported flap tip benefit of 3-4 dB was achieved [Ref. 169]. 

The current technology readiness level associated with flap tip fences is estimated to be TRL 8. 

However, due to their noise reduction spanning only a slight portion of the frequency spectrum, combined 

with the relative importance of flap noise when compared to other airframe noise sources, this technology 

was not considered for further analysis. 

B.54 Ejectors 

An ejector is a duct placed around the exhaust of a jet engine, circumferentially lined with passages, 

which allows the entrainment of ambient air into the main exhaust flow. By mixing the high-speed, hightemperature jet exhaust with the cooler, lower-speed ambient air, the resulting mixed flow velocity is 

decreased. As a result, the intensity of the turbulent mixing process between the primary hot stream and 

the secondary ambient stream is decreased, resulting in a reduction in noise generated by the exhaust jet 

[Ref. 170]. Depending on the ejector inlet geometry, the addition of this secondary air increases the static 

thrust by creating a suction on the inlet lip [Ref. 171]. It was reported by General Electric that the 

implementation of an ejector on the Northrop T-38 Talon resulted in a thrust improvement of 1-2% at 

takeoff conditions [Ref. 172]. Furthermore, flow mixing may be enhanced through the use of mixing 

lobes, vortex generators, etc. However, the implementation of these devices may themselves generate 

additional high frequency noise. This may be combated with liner technology along the interior surfaces 

of the ejector. In practical applications there can be a significant weight penalty associated with the 

implementation of an ejector. To maximize noise reduction, the length of the ejector must grow to such an 

extent that mixing becomes as complete as possible. However, longer ejectors add weight and frictional 

flow losses that contribute to an ultimate reduction in the thrust augmentation potential of the ejector. 

Depending on the relative diameters of the jet engine exhaust and the ejector inlet, ram drag effects could 

also become significant during flight conditions. 

Although the current TRL of this technology is estimated at 8, it was not considered for this study due 

to the estimated weight penalties associated with the technology. While benefits can be achieved during 

takeoff, in terms of noise, the weight associated with the ejector must be carried through the entire flight. 

B.55 Thrust Vectoring Open Rotor 

Thrust vectoring open rotor engine configurations can be achieved in two ways. The first is to thrust 

vector the jet exhaust, described in the Subsection B.71. The second is to enable rotation of the engine, 

which would allow for STOL, or even VSTOL, capabilities. This engine configuration would be similar 

to that found on the V-22. As engine sizing is driven largely by takeoff performance, VSTOL and STOL 

could potentially allow for decreased engine sizing. However, as described by Chana and Sullivan, during 

takeoff and landing, the stationary wing acts as a flat plate in the downwash of the tilted rotor, causing a 

download [Ref. 173]. This download reduces the net lift of the vehicle, thus limiting payload and range. 

In addition, tilt rotor vehicles show a decrease in cruise aerodynamics compared to traditional fixed 

propulsion aircraft. Figure B-3 illustrates the increase in parasite drag of the V-22 compared to 

commercial jet transports [Ref. 174]. Due to the increase in passengers, the jet transport aircraft have 

increased wetted areas. If a V-22-like vehicle was sized to equivalent passenger counts, the total wetted 

area and parasite area would be expected to increase even more. Figure B-3 illustrates the increase in 

cruise efficiency of the C-130, a fixed rotor configuration, compared to the V-22. In addition, Figure B-3 

shows the airspeed limitations of the V-22 tilt rotor configuration for optimum cruise performance. The 

current technology readiness level associated with open rotor thrust vectoring is estimated to be TRL 5; 

however, due to the increase in vehicle drag, and conventional high-lift systems proving sufficient, this 

technology was not investigated further. 

NASA/CR—2010-216798 134
Figure B-3. Tilt Rotor Aerodynamics 

B.56 Slat Tip Fences 

Slat tip fences are flat plate obstructions employed at the inboard and/or outboard portions of wing 

slats. They can be aerodynamically contoured depending on the platform on which they are utilized, 

extending below the lower surface of the slat. Their purpose is to prevent the shedding of streamwise 

vortices from the edges of the slat caused by the discontinuity in the wing surface upon slat deployment. 

These vortices have been associated with high levels of airframe noise during landing. This technology 

can be applied to any wing slat, and generally can be retrofitted to any wing slats that do not currently 

implement them. There will be a slight weight penalty due to their implementation. Failure of this 

technology would consist of failure of the attachment of the slat tip fence, at which point it would 

separate from the slat and fall from the aircraft. It could potentially cause damage to the leading edge of 

the main wing body, leading to a loss of the acoustic reduction benefit for which it is designed. In a recent 

experimental study, source noise reduction using an inboard slat tip fence was in the range of 4 dB, while 

at the outboard tip, a reduction of 2 dB was achieved [Ref. 153]. Weights associated with these devices 

are minimal. Although slat tip fences are currently estimated to be at a TRL of 5, because of the limited 

noise reduction potential of the technology, the incompatibility of slats with laminar flow, and the relative 

contribution of slat noise to the overall aircraft EPNL, this technology was discarded from further 

examination. 

B.57 Cryogenic Motors 

Employing motors cooled to superconducting temperatures would allow for operation at extremely 

high efficiencies. However, the engine architectures that would employ this type of motor are generally 

more complex than the conventional gas turbine cycle. Using motors for propulsive systems requires 

them to be driven by other devices, such as generators or fuel cells, and requires them to drive propellers 

to generate thrust. State-of-the-art advanced motors have specific power densities in the range of 2 hp/lb, 

meaning that in order to create enough takeoff power, the motor would have to be extremely heavy. In 

addition, a cryogenic motor necessitates the use of a cooling system to keep its temperature substantially 

low. This would generally take the form of liquid gas (e.g. hydrogen), requiring additional weight, 

insulation, and storage requirements. 

A cryogenic motor could operate with nearly no emissions, but at a current TRL estimated to be 1, it 

has been deemed impractical for the N+3 timeframe. 

B.58 Porous Plug Designs 

Porous nozzle plug designs consist of nozzle plugs which have a degree of porosity in the form of 

holes and a typically hollow interior. They are generally associated with the reduction of shock associated 

NASA/CR—2010-216798 135
noise in supersonic jets by controlling the formation of shocks and shock cells. However, when applied to 

subsonic flows, benefits are also achieved in noise reduction. This is mainly due to the reduction of shear 

noise. Subsonic noise suppression has been attributed to a reduction of the length of the jet potential core. 

It has also been proposed that the centerbody modifies the turbulent structure of the jet by inhibiting 

vortex pairing. This technology could be applied to any aircraft which uses a plug nozzle, by replacing the 

solid plug with a porous counterpart. In the case of an aircraft which does not have a plug nozzle 

implemented, the retrofitting of such a nozzle would be significantly more difficult. There would be a 

need to incorporate not only the plug nozzle, but the structure to maintain it in place. A porous plug 

nozzle design could suffer three modes of failure: a rupture of the structure of the plug due to its porosity 

(and possible weaker structural integrity), melting closed of some or all of the holes in the plug, or the 

detachment of the plug from the engine system. In the first case, it is likely that thrust losses will occur 

and noise would increase in certain frequencies, due to vortex shedding off of the rupture location. In the 

second case, with some or all holes melted closed, the plug would behave more like a solid plug with 

more noise associated with it. In the third mode of failure, the detachment of the entire plug, engine 

performance would suffer greatly due to the radically different geometric nozzle conditions compared to 

those for which it was designed. Aeroacoustic and thrust measurements were conducted by Bauer et al 

[Ref. 175]. In this study, eight different porous nozzle configurations were studied (mostly for supersonic 

applications, but at one subsonic pressure ratio) with various temperature core flows. OASPL reduction 

was measured at multiple directivity angles, and thrust losses were reported for each configuration. 

Weight penalty is assumed to be negligible, as porosity is introduced into a preexisting structure. 

Currently estimated at a TRL of 4, it was determined for this study that the thrust losses associated 

with the use of the plug nozzle outweighed the potential acoustic benefits in the subsonic flight regime. 

B.59 Porous Ceramic Materials 

Porous ceramic materials are being explored due to their lightweight nature and both their capacity 

for acoustic absorption and tolerance to high temperatures. By varying the porosity, pore size, flow 

resistance, thickness, and length, specific frequencies can be targeted depending on the location of their 

use in the engine core. The most common projected use for this material is as a nozzle or ejector liner 

where it can reduce noise associated with the jet engine exhaust while withstanding the heat produced by 

the combustor. In fact, porous ceramic materials can withstand temperatures in excess of 1000°C 

[Ref. 176]. 

The current TRL of this technology has been estimated to be 2; however, due to the implementation 

of variable nozzle geometry and the exclusion of ejector technology, this technology was not pursued 

further in the course of this study. 

B.60 Porous Flap Side Edges 

Streamwise vortex shedding from flaps is a contributor to airframe noise. These vortices have the 

effect of entraining the wake from the main wing element, and with the pressure fluctuations from the 

combined flow impinging on the side edge of the flap, whereby airframe noise is generated. By 

incorporating a flap side edge with porosity, the turbulent stresses contained in this vortex are reduced 

and the vortex becomes displaced further from the flap surface. In this manner, the noise from the 

impingement on the flap surface is minimized. This technology can be applied to any wing flap in which 

there is space to include them, and generally could be retrofitted to any wing flaps that do not currently 

implement them. There will be a slight weight penalty due to their use. Failure of this technology would 

consist of failure of the attachment of the porous flap side edge, at which point it would separate from the 

flap and fall from the aircraft. This would lead to a loss of the acoustic reduction benefit for which it was 

designed. As reported in Fink and Bailery, flap side edge porous treatment led to a total flap noise 

reduction of up to 2 to 3 dB [Ref. 177]. In another experimental study, it is reported that the aerodynamic 

of porous flap side edges is negligible [Ref. 178]. 

NASA/CR—2010-216798 136
With a currently estimated TRL of 4, it was decided to not pursue this technology further due to its 

limited capacity for noise reduction in a narrow frequency range, combined with the negligible 

contribution of flap noise to the overall EPNL of the system. 

B.61 Slat Trailing Edge Serrations 

Trailing-edge serrations on wing slats involve cutting or constructing V-shaped grooves into the slat 

material. In the presence of serrations, the periodic vortex shedding from the trailing edge of the slat that 

would typically occur becomes less coherent in the spanwise direction. In this manner, the noise radiated 

from the trailing edge of the slat is reduced. This technology can be used on the trailing edge of any slat. 

Depending on the construction of a particular slat, retrofitting the technology may be possible. Because 

no additional material must be introduced into the slat to implement the serrations, there will be a 

negligible weight penalty. This technology is built into the slat of the wing, and as such the only available 

mode of failure would be for serrations to break apart or bend in the flow. If this were to happen, then the 

aerodynamic modifications to the flow associated with the presence of the serration would be nullified or 

changed, and localized pockets of higher noise generation would be created. In an experimental wind 

tunnel study, a peak noise reduction of 3 dB was achieved by using slat trailing edge serrations 

[Ref. 153]. However, according to this study, the average noise reduction across the frequency range was 

approximately 2 dB. Weight penalties are negligible for this technology. However, due to the limited 

noise reduction potential of this technology, combined with the obsolescence of the technology in the 

presence of laminar flow (which requires complete removal of leading edge slats), this technology has 

been discarded for further examination. 

With a currently estimated TRL of 4, it was decided to not pursue this technology further due to its 

limited capacity for noise reduction combined with the removal of slats associated with the integration of 

laminar flow. 

B.62 Porous Slat Pressure Surfaces 

Porous slat pressure surfaces are designed to reduce slat noise. Based on wind tunnel results 

performed by NASA, U.S. partners, and the University of Southampton, vortices generated from the slat 

trailing edge were found to be a significant portion of the slat noise. These vortices are often amplified in 

the region between the slat and the main element through resonant reflection. Porous slat pressure 

surfaces have been found effective in reducing slat trailing edge vortices strength and when placed 

appropriately, can be effective in reducing flap noise by weakening the vortex generated from the flap 

side edge. Porous pressure surfaces reduce this vortex by reducing the intensity of pressure fluctuations 

near the trailing edge of the slat. According to Khorrami and Choudhari, porous slat pressure surfaces can 

reduce slat far field noise by more than 20 dB [Ref. 179]. In addition to the SPL level reduction, this 

technology increases the dominant frequency approximately 25%. However, this dominant frequency on 

full scale applications is expected to exceed human auditory range, thus this shift in frequency would lead 

to a further reduction in noise. Porous slat pressure surfaces are surfaces composed of a finely woven wire 

mesh and could use material that is currently used for turbine blade wall cooling or acoustic treatment of 

engine ducts. This use of commonly used material would only be effective if designed and applied 

correctly. Weight penalties are negligible for this technology. 

Currently estimated at TRL 4, this technology was not investigated further due to the obsolescence of 

the technology in the presence of laminar flow (which requires complete removal of leading edge slats). 

B.63 Micro-Tab Features 

Micro-tab features consist of small retractable tabs embedded inside the flap of a high-lift wing that 

are deployed on either the upper or lower surface of the flap perpendicular to the oncoming flow in order 

to increase the effective camber of the wing. The height of these tabs is on the order of boundary layer 

NASA/CR—2010-216798 137
thickness, so they do not result in excessive drag penalties. This increase in lift on the wing from the 

micro tabs allows for the flap and slat setting to be slight eased leading to reduction in flap and slat 

generated noise. If the micro-tabs are designed to be retractable, complexity would be added through 

wiring and controls. The weight penalty of the micro-tabs themselves is negligible. Noise in high-lift 

systems is extremely configuration dependant and any increase in flap angle will lead to an increase in 

noise as a result of increasing the load on the high-lift components and may generate stronger vortex 

shedding and turbulent wake downstream. In an experiment by Kuo and Sarigul-Klijn, a baseline airfoil 

with 30 degree slat and flap deflection at an angle of attack of 6 degrees with no noise reduction 

technologies was used to compare against an airfoil with a 20 degree deflected slat and flap deflection at 

an angle of attack of 8.8 degrees with microtab integration [Ref. 180]. Both high lift airfoils were derived 

from a Boeing 737 type midspan airfoil with a single flap. It was found that beyond 100 Hz, the micro tab 

configuration reduced wing noise levels by 2-5 dB over the entire frequency spectrum compared to the 

baseline configuration. When the A-weighted acoustic results are compared, it can be seen in Figure B-4 

that the microtab configuration shares a noise spike at a frequency of approximately 315 Hz and a 

reduced noise spike at 630 Hz when compared to the baseline configuration. The noise spike of the 

baseline at 1260 Hz is not present in the microtab configuration. Microtabs were found to reduce the 

overall noise levels by 2.3 dB. In addition, the microtab configuration was found to decrease drag by 5%. 

Figure B-4. Noise Attenuation With and Without Microtab Features 

Currently estimated at TRL 4, this technology was not investigated for further analysis due to the 

relative importance of the wing noise when compared to other airframe noise sources in addition to 

decreased effectiveness in the presence of laminar flow (which requires complete removal of leading edge 

slats). Fuel burn by mission segment for preferred configuration. 

B.64 Fiber Metal Laminates 

Fiber metal laminate (FML) is a class of metallic materials consisting of a laminate of several thin 

metal layers bonded with layers of composite material, commonly glass fiber and aluminum that was 

designed to enhance the fatigue performance of aluminum structures. FMLs combine the high bearing 

strength and impact properties of metals with the high fatigue, strength, and stiffness properties of 

composite matrices [Ref. 181]. This allows the material to behave much as a simple structure with 

specifically tailored advantages regarding structural properties including crack propagation. 

This increase in fatigue and other material properties, offers potential weight savings. This technology 

has not seen extensive military use, but has been utilized in cargo doors and fuselage skins of nonmilitary 

vehicles. Based on internal research, FMLs have shown weight savings for only very few applications. 

Composite structures would be lighter than FML for tension dominated structures. In addition, 

NASA/CR—2010-216798 138
compression strength dominated structures would gain increased weight savings from composites versus 

FML. FMLs have been forecasted to be not competitive for use with min-gauge structures, which are the 

majority of structures found in flying wing bodies. Potential acreage applications on vehicles would be 

impact prone areas including cargo floors, ramps, and doors. 

Although positive benefits are expected due to increased fatigue and impact properties, internal 

analysis has revealed that FML effects with regard to the overall vehicle weight is minimal when weighed 

together with the relative immaturity of the technology. 

Currently estimated at TRL 4, this technology was not investigated further due to having no 

significant weight benefits. 

B.65 Metallic Digital Direct Manufacturing 

Metallic Digital Direct Manufacturing (MDDM) is a manufacturing process that creates physical 

parts from a 3-D CAD file or data using additive fabrication techniques [Ref. 182]. Several MDDM 

techniques use a laser or electron beam as the main energy source. The system aims to create a 3-D part 

with the sequential solidification of 2-D layers. The process uses one of three methods for injecting 

material into the energy source, which fuses the material onto the surface, creating a layer. This process is 

repeated until the part is completed. This allows for the fabrication of extremely complex parts with 100% 

density. Manufacturing speed, preciseness, and productivity is also increased. 

MDDM can be used for any metallic component in the vehicle including titanium. Through internal 

research, it was found that MDDM would offer potential weight savings through manufacturing of 

complex integrated part that would normally be comprised of fastened sub-components, reducing part 

count. Although positive benefits are expected due reduction in part count, internal analysis has revealed 

that MDDM effects with regard to the overall vehicle weight is minimal when weighed together with the 

relative immaturity of the technology. 

Currently estimated at TRL 4, this technology was not investigated further due to having no 

significant weight benefits. 

B.66 Open-Celled and Stamped Composite Structures 

Open-celled and stamped composite structures use a composite-based trussed beaded or sinewave 

structure. Through the use of composites and structural layout, structural weight reduction is possible. 

Open-celled and beaded structures offer high volumetric efficiency, high dynamic damping, mechanical 

isotropy, increased shelf life and ease of manufacturing. The current baseline is composed of nonbuckled 

metallic and composite structures, with sandwich structures as the leading candidate. Based on internal 

research and testing, depending on application and on the minimum gage definition for composite 

sandwich, open-celled and stamped composite structures could offer a part weight savings up to 20%. 

Candidate applications for open-celled structures are minimum gage structures with low loading such as a 

trussed substructure’s webs with holes. Candidate applications for stamped structures, such as beaded or 

sinewave structures, are structures that have some degree of loading and hence thickness. However, this 

application would be limited to substructural webs. 

Reduced vehicle weight would reduce fuel burn and field length for landing and takeoff. When the 

relative system level weight reduction from this technology was compared to other airframe technologies, 

it was felt that an additional structural technology with similar effects was not warranted as its model 

would fall in line with the result of the others. 

The current technology readiness level associated with open-celled and stamped composites is 

estimated to be TRL 5; however, due to the weight reduction being applicable to select applications and 

larger scale weight reduction being in line with other structural technologies, this technology was not 

investigated further. 

NASA/CR—2010-216798 139
B.67 Post-Buckled Composite Structures 

Post-buckled composite structures deviate from the traditional design of allowing initial buckling to 

occur at or slightly above the ultimate load of a structure and lead to reduced weight and manufacturing 

costs [Ref. 183]. Post-buckled structures demonstrate that certain structures exhibit significant strength 

beyond that demonstrated at initial buckling [Ref. 184,185]. An increased reliance on design and model 

analysis is required due to the nonlinear material and geometric behavior associated with post-buckled 

composite structures. This structural design allows for the use of minimum gage designs in candidate 

components. Aggressive use of minimum gage designs would result in a major part weight reduction in 

structures that are not extremely heavily loaded. The current baseline is nonbuckled composed of 

nonbuckled metallic and composite structures with sandwich as the leading candidate. Based on internal 

research, depending on the minimum gage definition for composite sandwich, post-buckled composite 

panels could result in parte weight reduction up to 20%. Candidate applications for post-buckled 

structures are structures that have some degree of loading and hence thickness and in both substructures 

and outer mold line skins. Large bulkhead subsections could potentially be post-buckled to reduce weight. 

Regarding outer mold line skins, many transport fuselage and deep section skins are mostly minimum 

gauge. Wing skins demonstrate potentially large benefits from post-buckled structures. Reduced vehicle 

weight would reduce fuel burn and field length for landing and takeoff. When the relative system level 

weight reduction from this technology was compared to other airframe technologies, it was felt that an 

additional structural technology with similar effects was not warranted as its model would fall in line with 

the result of the others. 

The current technology readiness level associated with post-buckled composite structures is estimated 

to be TRL 5; however, due to the weight reduction being applicable to select applications and larger scale 

weight reduction being in line with other structural technologies, this technology was not investigated 

further. 

B.68 Direct Manufacturing/Laser Sintering of Polymeric Materials 

Direct part manufacturing/laser sintering is an additive process capability that uses net-shape and 

layer build technology. Components are produced from a machine without tooling with engineering data 

generated from a 3-D computer model. The component is produced to be dimensionally accurate and 

requires little to no postprocessing or sizing prior to finishing. The machine uses a laser to machine away 

excess ceramic material and fuses together the polymer that will be used for the part. 

The parts produced with this manufacturing method are limited to non-load bearing structures within 

subsystem assemblies and subassemblies. This technology is primarily used to reduce part count, tooling, 

post processing, part weight, and is able to fabricate complex designs. Nylon 12 is currently used with 

this manufacturing technique in the Global Hawk and in the F-18. However, Nylon 12 is limited to 

temperatures less than 180 degrees Fahrenheit. Higher temperature materials are needed in order to see a 

more significant vehicle level weight reduction. 

The TRL of direct manufacturing/laser-sintering is 9 for Nylon 12 and 3 for higher temperature 

materials. Due to the low weight reduction relative to other technologies on the vehicle level, this 

technology was not investigated further. 

B.69 Slat Gap Liners 

This technology entails lining the lower surface of the slat and portions of the main wing leading edge 

with acoustic liners to suppress noise in certain frequency ranges, particularly those associated with the 

convection of turbulence through the slat gap and past the sharp trailing edge of the slat. Due to the 

embeddable nature of acoustic liners, there is little to no effect on the aerodynamic performance of the 

high-lift device caused by the presence of the liner, and they are easily tunable to particular frequencies in 

the design phase. However, current liner technology is passive in nature, so it is possible that this 

NASA/CR—2010-216798 140
technology would not be optimized for all conditions. Depending on the attachment method, slat gap 

liners could be affixed to any wing that employs slats for added lift capabilities. This will adversely 

(though minimally) affect the aircraft weight, though due its passive nature there will be no additional 

requirements in terms of electrical cables or hydraulics. Failure modes of the technology depend on how 

the liners are affixed or embedded into the slat and main wing leading edge. If it is a retrofit situation 

where the attachment method could fail, then the liner sections could become projectiles in the case of a 

failure. However, due to their size and weight, any capacity for damage would be minimal. In Smith et al, 

it is shown through wind tunnel experimentation that in the range of 2-10 kHz, a fully-lined (slat cove and 

main wing leading edge) high-lift device averages around 1.5 dB noise reduction averaged over 

directivity angles [Ref. 186]. This test also confirmed that a properly attached liner has no notable effect 

on the aerodynamic performance of the wing through a comparison of CL-α curves with and without the 

liner installed. Weight could effectively be reduced by the use of liners since they are porous in nature, 

though to accommodate for attachment requirements, it will be assumed that these benefits and penalties 

balance out. However, due to the limited noise reduction potential of these liners, combined with the 

obsolescence of the technology in the presence of laminar flow (which requires complete removal of 

leading edge slats), this technology has been discarded for further examination. 

The TRL of slat gap liners is currently estimated to be 3. Due to its limited capacity for noise 

reduction combined with the removal of slats associated with the integration of laminar flow, slat gap 

liners were not considered for further investigation. 

B.70 Streamlined Slat Tracks 

Current slat tracks guide the slat between its deployed and undeployed state with little consideration 

given to the noise and aerodynamic penalties that may exist. These components generate noise, and highlift noise on approach is generally dominated by the slat. Streamlining the slat tracks is a potential 

contributor to high-lift noise reduction. Slat tracks could either be modified using add-on treatments to fill 

cavities and generally smooth out the surfaces, or could be redesigned at an earlier stage in the aircraft 

design process to give a more aerodynamic geometry. In the case of add-on treatments, a small degree of 

weight would be added to the system depending on the material used in the treatment. For newly designed 

streamlined slat tracks, the design could represent a reduced component weight. 

In the case of an add-on treatment for streamlining the slat track, this treatment could become 

loosened or damaged with use. This may affect the retraction of the slat, leading to undesirable 

performance characteristics after the lift-off sequence. If the slat track were geometrically redesigned with 

no add-on parts, failure would be more dramatic (but less likely). In this case a possible mode of failure 

would be the structural weakening or breaking of the slat track, which would have serious consequences 

for the retraction of the slat. 

According to Dobrzynski et al [Ref. 187], slat tracks constitute sources of excessive flow noise, on 

the order of 8 dB higher than the rest of the slat. A flight test of an add-on slat track treatment showed a 

reduction of 0.4 EPNdB due mainly to noise reduction at and above the 315 Hz 1/3-octave band 

[Ref. 188]. This work also reports that no changes in drag performance was noticed or expected as a result 

of the add-on treatments. 

Although positive benefits are expected due to the implementation of streamlined slat tracks, 

experimental measurements have revealed that their effects with regard to the overall EPNL of the aircraft 

is minimal when weighed together with the relative immaturity of the technology. Currently estimated at 

TRL 3, this technology was not investigated further. 

B.71 Thrust Vectoring Nozzles 

Thrust vectoring is a technology that allows for the vectoring of engine thrust improving 

maneuverability and allowing STOL capabilities. According to Friehmelt, vehicles that were integrated 

NASA/CR—2010-216798 141
with thrust vectoring nozzles demonstrated a 38% shorter take-off distance and 60% shorter landing roll 

[Ref. 189]. However, Friehemlt states that up to 30% of the net thrust can be lost due to undesirable 

deflection in portions of the flow. This decrease in thrust efficiency would increase fuel burned 

throughout the vehicle mission. According to other studies, mechanical thrust vectoring systems result in 

significant increased weight and present integration issues [Ref. 190]. Mechanical thrust vectoring 

systems are present on many modern military vehicles. Fluidic thrust vectoring offers improved thrust 

efficiency with a less significant weight penalty. However, this technology is still in early stages of 

development. 

Figure B-2 compares the lift capabilities of thrust vectoring and flap configuration to a conventional 

flap configuration. Vectored thrust with flaps shows considerable improvements for STOL capabilities 

compared to a conventional flap configuration. 

Although positive field length benefits are expected due to the implementation of thrust vectoring, the 

structural and fuel weight penalties associated with both fluidic and mechanical thrust vectoring outweigh 

the field length improvements for the given mission. As a result of integrated propulsion and advanced 

field length configurations being dismissed in Section 7.2.9, and low TRL level of fluidic thrust 

vectoring, thrust vectoring was not investigated further. 

B.72 Metal Foam Liner 

Traditional DOF liners are located in the inlet where metal foam liners allow for the placement closer 

to the engine rotor, which is closer to the noise source and can provide increased effectiveness in noise 

reduction. Metal foam liners would be able to reduce engine far field noise by providing a pressure relief 

region and altering the acoustic near field. This can reduce noise generation sources and allow improved 

reduction in regions further away from the rotor. Based on the work of Sutliff and Jones Ref. 191, it was 

found that if a foam-metal liner was installed over the rotor, a possible noise reduction of 4 dB was 

attainable. 

A tradeoff exists between the absorption capacity of the liner and the drag that is generated. 

Additionally, the weight and volume penalties associated with acoustic liners increases as the thickness 

and area increases. 

Metal foam acoustic liners are currently estimated to be TRL 6, this technology was not considered 

for further investigation due a two DOF liner proving to be superior for the current configuration. 

NASA/CR—2010-216798 142
Appendix C 

Preferred Vehicle Aerodynamics

NASA/CR—2010-216798 143
Appendix C Preferred Vehicle Aerodynamics 

Tables C-1 through C-2 and Figures C-1 through C-3 show preferred vehicle aerodynamics 

Table C-1. Preferred Vehicle High Speed Drag Polar at 45,000 feet Altitude 

Table C-2. Preferred Vehicle Reynold’s Number Drag Coefficient Corrections 

Figure C-1. Preferred Vehicle Low-Speed Drag Polar for Takeoff at 0 feet Altitude 

NASA/CR—2010-216798 144
Figure C-2. Preferred Vehicle Low-Speed Drag Polar for Landing at 0 feet Altitude 

Figure C-3. Preferred Vehicle High-Speed Drag Polar, Rn associated with 45,000 feet Altitude 

NASA/CR—2010-216798 145
Appendix D 

Vehicle OASPL Maps 

NASA/CR—2010-216798 146
Appendix D Vehicle OASPL Maps 

Figure D-1 shows an example overall sound pressure level (OASPL) result for a random vehicle 

source. OASPL is plotted against polar angle, from -90 to +90 degrees, measured about the fuselage from 

one wing to the other, and against the azimuth angle, from 0 to 180 degrees, measured in the plane of the 

aircraft from the nose to its tail. Figures D-2 through D-9 use these standardized axes and scales, allowing 

for easy, uniform comparison. 

Figure D-1. OASPL Example Source Plot 

NASA/CR—2010-216798 147
D.1 Reference Vehicle – OASPL 

Figure D-2. Noise Source Levels Reference Vehicle at 250 feet Altitude [Descent-Landing] 

Figure D-3. Noise Source Levels Reference Vehicle at 250 feet Altitude [Takeoff-Climb] 

NASA/CR—2010-216798 148
D.2 Advanced Tube-and-Wing, Scaled CFM56-3B1, No Technologies – OASPL 

Figure D-4. Noise Source Levels ATW, Scaled CFM56-3B1, No Advanced Technologies 

at 250 feet Altitude [Descent-Landing] 

Figure D-5. Noise Source Levels ATW, Scaled CFM56-3B1, No Advanced Technologies 

at 250 feet Altitude [Takeoff-Climb] 

NASA/CR—2010-216798 149
D.3 Hybrid Wing-Body, Scaled CFM56-3B1, No Technologies – OASPL 

Figure D-6. Noise Source Levels HWB, Scaled CFM56-3B1, No Advanced Technologies 

at 250 feet Altitude [Descent-Landing] 

Figure D-7. Noise Source Levels HWB, Scaled CFM56-3B1, No Advanced Technologies 

at 250 feet Altitude [Takeoff-Climb] 

NASA/CR—2010-216798 150
D.4 Preferred Configuration – OASPL 

Figure D-8. Noise Source Levels Preferred Configuration at 250 feet Altitude [Descent-Landing] 

Figure D-9. Noise Source Levels Preferred Configuration at 250 feet Altitude [Takeoff-Climb] 

NASA/CR—2010-216798 151
Appendix E 

QFD Spreadsheet 

NASA/CR—2010-216798 152
Appendix E QFD Spreadsheet 

The QFD spreadsheet is shown in Figure E-1. 

Figure E-1. QFD Worksheet (Sheet 1 of 6) 

NASA/CR—2010-216798 153
Figure E-1. QFD Worksheet (Sheet 2 of 6) 

5 Airframe Concept #1.05 3-D Woven / Stitched Composites 

Advanced 3-dimensionally woven or stitched fiber composites to enable integrated bonded structures with standard and intermediate modulus carbon fibers 5 19.52 5 7.81 5 19.52 3 6.00 52.8 3 34.2 1.000 1 1 52.8

7 Airframe Concept #1.07 Light-Weight Foamed Composites

Single process fabrication of foamed composite core with skins 5 19.52 5 7.81 5 19.52 3 6.00 52.8 3 34.2 1.000 1 2 52.8

9 Airframe Concept #1.09

Integrated Aeroservoelastic (ASE) Structures

Structural weight reduction through gust/flutter/maneuver load control, potential reduction of actuation system weight through active aero-elastic wing technology 5 19.52 5 7.81 5 19.52 3 6.00 52.8 3 34.2 1.000 0 0 52.8

20 Aero Concept #2.06

Distributed Exhaust Nozzle Flap (DEN Flap)

Distributed Exhaust Nozzles Integrated Into Flap Provide Flow Control/Supercirculation. 1 3.92 9 13.33 5 19.52 7 17.50 54.3 2 46.1 0.961 2,3,4 13 52.1

22 Aero Concept #2.08

Unsteady Circulation Control (Offtake Engine Flow Used for Circulation Control via Blown Flap, IBF, Coanda Jet Flap, etc.)

Pulsed Effectors Provide Similar Benefits as Steady CCW 1 3.92 9 13.33 5 19.52 7 17.50 54.3 2 46.1 0.961 2,3,4 0 52.1

101 Prop Concept #4.33 Staged Combustors

Series of combustion chambers burn leaner so that total NOx production is lowered, and slight thermal efficiency gains realized. 3 10.00 3 4.00 7 29.16 3 6.00 49.2 6 18.3 1.053 0 0 51.8

100 Prop Concept #4.32 Water Injection Reduced emissions. 3 10.00 5 7.81 5 19.52 5 11.71 49.0 6 18.3 1.053 0 0 51.6

93 Prop Concept #4.25 Variable Pitch Fan Blades

Vary blade pitch to optimize FPR at various conditions, eliminates need for thrust reverser, reduce windmilling drag. 3 10.00 5 7.81 5 19.52 5 11.71 49.0 5 22.2 1.040 1,4 -1 51.0

72 Prop Concept #4.04 Swept Fan Outlet Guide Vanes

Decrease pressure losses especially for lower FPR designs. Reduce fan-tip noise. 5 19.52 5 7.81 3 10.00 5 11.71 49.0 5 22.2 1.040 3,4 22 51.0

23 Aero Concept #2.09 Swept-Wing Laminar Flow 

Reduced Skin Friction by Increasing Regions of Laminar Flow Through Shaping 7 29.16 3 4.00 3 10.00 3 6.00 49.2 4 27.2 1.023 2,3 6 50.3

90 Prop Concept #4.22 Fuel Cell APU

Replace traditional APU with regenerative fuel cells. 5 19.52 3 4.00 5 19.52 3 6.00 49.0 4 27.2 1.023 2,4 20 50.2

NASA/CR—2010-216798 154
Figure E-1. QFD Worksheet (Sheet 3 of 6) 

90 Prop Concept #4.22 Fuel Cell APU fuel cells. 5 19.52 3 4.00 5 19.52 3 6.00 49.0 4 27.2 1.023 2,4 20 50.2

92 Prop Concept #4.24 TO/C-Assist Propulsion Unit

Detachable takeoff/climb autonomous propulsion units propel vehicle to cruise altitude, then return to airport increasing cruise efficiency. 5 19.52 9 13.33 5 19.52 5 11.71 64.1 1 100.0 0.783 2,3,4 31 50.2

24 Aero Concept #2.10

Microvortex Generators and Riblets Reduced Turbulent Skin Friction Drag· 7 29.16 3 4.00 3 10.00 3 6.00 49.2 3 34.2 1.000 1,2 4 49.2

69 Prop Concept #4.01

Conformal Polymer Solar Cell Power Augmentation

Quantum dot nanoparticles embedded in polymer capture extended EM spectrum beyond visible to infrared. 5 19.52 3 4.00 5 19.52 3 6.00 49.0 3 34.2 1.000 2,4 -1.00 49.0

95 Prop Concept #4.27 Intercooled Compressor Stages

Increase flow density for higher OPR, more efficient combustion, and decreased flowpath sizing. 5 19.52 3 4.00 5 19.52 3 6.00 49.0 3 34.2 1.000 4 25 49.0

91 Prop Concept #4.23 Hybrid All-Electric Propulsion

Combine multiple electric-power generating options, including isotope power cells or heat exchangers, chemical batteries, fuel cells, ultracapacitors, and solar. 1 3.92 1 1.57 9 33.33 5 11.71 50.5 2 46.1 0.961 2,3,4 58 48.6

30 Noise Concept #3.02

Shielding (Fan, Jet, other Propulsive Sources)

Propulsion Noise Sources Located Over Wings, Fuselages, and Tails to Impeded Acoustic Propagation 3 10.00 3 4.00 3 10.00 9 20.00 44.0 9 10.0 1.080 3,4 15 47.5

42 Noise Concept #3.14 Moveable Winglets

Minimized Vortex Shedding for a Broader Envelope of Operation Points 3 10.00 3 4.00 7 29.16 3 6.00 49.2 2 46.1 0.961 2 0 47.2

58 Noise Concept #3.30 Landing Gear Assembly Fairings Streamlined Gear Assembly 3 10.00 3 4.00 3 10.00 9 20.00 44.0 8 12.3 1.072 0 0 47.2

59 Noise Concept #3.31

Landing Gear Assembly Component Integration

Reduction of Turbulence Around Gear Assemblies 3 10.00 3 4.00 3 10.00 9 20.00 44.0 8 12.3 1.072 0 0 47.2

88 Prop Concept #4.20 Combined/Compound Cycles

Integrate multiple thermodynamic exothermic cycles into a single "core" engine - Otto, Brayton, Sterling, thermoacoustic cycles, etc. 5 19.52 3 4.00 5 19.52 3 6.00 49.0 2 46.1 0.961 4 16 47.1

33 Noise Concept #3.05 Fan Blade Sweep Design Reduced Fan Tip Speed 3 10.00 3 4.00 3 10.00 7 17.50 41.5 8 12.3 1.072 1,4 20 44.5

41 Noise Concept #3.13 Curved, Scarfed Inlet Ducts Self Shielding Inlet Ducts 3 10.00 3 4.00 3 10.00 7 17.50 41.5 8 12.3 1.072 3,4 51 44.581 Prop Concept #4.13 Compressor Flow Control

Active blade secondary flow injection may improve surge margin and decrease stage count and/or blade count. 5 19.52 5 7.81 3 10.00 3 6.00 43.3 4 27.2 1.023 4 6 44.3

15 Aero Concept #2.01 Steady Fluidic Separation Control

Blowing/Suction for Separation Control Provides ΔCL Improvement for STOL 1 3.92 9 13.33 5 19.52 3 6.00 42.8 4 27.2 1.023 2 7 43.8

6 Airframe Concept #1.06

M5 Fiber Ultra High Performance Fiber

Advanced organic fiber produced by Magellan Systems International in partnership with DuPont Advanced Fiber Systems 5 19.52 5 7.81 3 10.00 3 6.00 43.3 3 34.2 1.000 1 0 43.3

16 Aero Concept #2.02 Synthetic Jet Separation Control

Synthetic Jet Provides Flow Control With Zero Net Mass Flow Jet for ΔCL Improvement 1 3.92 9 13.33 5 19.52 3 6.00 42.8 3 34.2 1.000 2 4 42.8

19 Aero Concept #2.05

Lower Surface Exhaust Nozzle and Flap

High-Cm Exhaust on Slotted Flap Keeps Flow Attached at Large Flap Deflections and Provides Supercirculation 1 3.92 9 13.33 5 19.52 3 6.00 42.8 3 34.2 1.000 2,4 1 42.8

NASA/CR—2010-216798 155
Figure E-1. QFD Worksheet (Sheet 4 of 6) 

34 Noise Concept #3.06 Fan Blade Flow Control PROVIDE DESCRIPTION - KYLE 3 10.00 3 4.00 3 10.00 7 17.50 41.5 4 27.2 1.023 1,2,4 15 42.568 Noise Concept #3.40 Nanotube Liner GTRI Developed 3 10.00 3 4.00 3 10.00 9 20.00 44.0 2 46.1 0.961 3,4 8 42.328 Aero Concept #2.14 Channel Wing Increased lift 1 3.92 7 11.66 5 19.52 3 6.00 41.1 4 27.2 1.023 2,3,4 7 42.171 Prop Concept #4.03 Plane Jet Combustion

Low emissions from better atomization, lower residence time. 3 10.00 3 4.00 5 19.52 3 6.00 39.5 5 22.2 1.040 4 -1 41.1

74 Prop Concept #4.06 Electric Subsystems

Electric fuel and oil pumps, provide mechanical efficiency over hydraulic equipment. 5 19.52 3 4.00 3 10.00 3 6.00 39.5 5 22.2 1.040 0 1 41.1

82 Prop Concept #4.14 Third Engine Stream

Additional flowpath with variable geometry will allow higher thrust levels at takeoff/climb and maintain higher cruise efficiency. 5 19.52 5 7.81 3 10.00 1 2.35 39.7 4 27.2 1.023 3,4 48 40.6

80 Prop Concept #4.12 Alternative Hydrocarbon Biofuels

Low VOC's produce much fewer emissions, and may produce carbon-neutral exhausts from algae-based fuels. 3 10.00 3 4.00 5 19.52 3 6.00 39.5 4 27.2 1.023 4 -2 40.4

77 Prop Concept #4.09 Embedded/Conformal Inlets

Compact inlet integrations may require less structure, and decrease wetted area. 5 19.52 3 4.00 3 10.00 3 6.00 39.5 4 27.2 1.023 2,3,4 65 40.4

87 Prop Concept #4.19

Active Compressor Clearance Control

Provide higher compressor efficiencies by minimizing blade tip losses by maintaining clearances. 5 19.52 3 4.00 3 10.00 3 6.00 39.5 4 27.2 1.023 4 9 40.4

97 Prop Concept #4.29 Low-Noise Fan Blades

Blade shaping (RESOUND) can reduce acoustic levels while retaining component efficiency. 3 10.00 3 4.00 3 10.00 7 17.50 41.5 2 46.1 0.961 3,4 21 39.9

83 Prop Concept #4.15 SLOTIP Fan & Fan Cowl Reduced structure weight & fan duct length. 5 19.52 5 7.81 3 10.00 1 2.35 39.7 3 34.2 1.000 3,4 21 39.7

8 Airframe Concept #1.08

Conformal Load-bearing Antenna Structures

Weight reductions through increased structural efficiency (no cutouts for bolt-on antennas), drag reduction through elimination of blades and antenna protuberance 5 19.52 3 4.00 3 10.00 3 6.00 39.5 3 34.2 1.000 0 0 39.5

12 Airframe Concept #1.12

Affordable Airframe Large Integrated Structures

Structural & Material Concepts For Affordable Yet Robust & Efficient Airframe Structures For Improved Persistence, Range, And Payload 5 19.52 3 4.00 3 10.00 3 6.00 39.5 3 34.2 1.000 1 1 39.5

99 Prop Concept #4.31

Ceramic Matrix Composite Turbine Blades

Reduce weight and improve thermal handling for hot section. 5 19.52 3 4.00 3 10.00 3 6.00 39.5 3 34.2 1.000 4 0 39.5

29 Noise Concept #3.01 DEN (Distributed Exhaust Nozzle)

Treated Engine Core Exhaust With SmallDiameter Nozzles Shifts Noise to High Frequencies 3 10.00 3 4.00 1 3.92 9 20.00 37.9 5 22.2 1.040 2,3,4 5 39.4

17 Aero Concept #2.03

Upper Surface Exhaust Nozzle and Flap

High-Cm Exhaust Keeps Flow Attached at Large Flap Deflections and Provides Supercirculation. Raised Nozzle Jet Flow Directed Onto Wing Upper Surface During High-Lift Mode 1 3.92 9 13.33 5 19.52 1 2.35 39.1 3 34.2 1.000 2,4 0 39.1

18 Aero Concept #2.04

Fluidic Thrust Vectoring for Upper Surface Exhaust Nozzle and Flap

Same as Upper Surface Exhaust Nozzle and Flap but Force Upper Surface Flow Towards Wing by Injecting Secondary Flow·Simplified Mechanisms 1 3.92 9 13.33 5 19.52 1 2.35 39.1 3 34.2 1.000 2,4 3 39.1

64 Noise Concept #3.36 Single DOF Liner 3 10.00 3 4.00 3 10.00 5 11.71 35.7 9 10.0 1.080 3,4 8 38.6

25 Aero Concept #2.11 Active Drag Reduction 

Reduced Skin Friction and Profile Drag by Using Microadaptive Flow Control, etc. 7 29.16 3 4.00 3 10.00 3 6.00 49.2 1 100.0 0.783 1,2 8 38.5

NASA/CR—2010-216798 156
Figure E-1. QFD Worksheet (Sheet 5 of 6) 

p g g p

66 Noise Concept #3.38 Multi DOF Liner 1 3.92 1 1.57 3 10.00 9 20.00 35.5 8 12.3 1.072 3,4 13 38.175 Prop Concept #4.07 Oil-Less Engine 5 19.52 3 4.00 3 10.00 3 6.00 39.5 2 46.1 0.961 0 3 38.076 Prop Concept #4.08 BLI Inlets

Accelerating streamtube creating airframe drag may provide higher propulsive efficiency independent of bypass ratio. 5 19.52 3 4.00 3 10.00 3 6.00 39.5 2 46.1 0.961 2,3,4 49 38.0

96 Prop Concept #4.28 Porous Ceramic Materials

Retain bulk-acoustic absorption properties with reduced weight. 5 19.52 3 4.00 3 10.00 3 6.00 39.5 2 46.1 0.961 1,3,4 15 38.0

84 Prop Concept #4.16

Low Tip Speed Counter-Rotating Ducted Fan

Lower tip speed increases efficiency, reduces weight, and will increase propulsive efficiency. 1 3.92 3 4.00 3 10.00 9 20.00 37.9 3 34.2 1.000 3,4 0 37.9

60 Noise Concept #3.32 Wheel Caps Elimination of Wheel Cavities 3 10.00 3 4.00 3 10.00 5 11.71 35.7 6 18.3 1.053 0 0 37.6

61 Noise Concept #3.33 Wheel Gap Filler Streamlined Wheel Assembly 3 10.00 3 4.00 3 10.00 5 11.71 35.7 6 18.3 1.053 0 0 37.638 Noise Concept #3.10

Aerothermodynamic Concepts (Inverted Flow, Thermal-Acoustic Shield)

Impedence Changes Inducing Shielding Effects 3 10.00 3 4.00 3 10.00 5 11.71 35.7 4 27.2 1.023 3,4 8 36.5

40 Noise Concept #3.12 Acoustic Excitation

Acoustically Excited Jets Go Violent When Excited by Pure Tones Which Increases Turbulence and Enhances Mixing 3 10.00 3 4.00 3 10.00 5 11.71 35.7 4 27.2 1.023 3,4 15 36.5

44 Noise Concept #3.16 Brush-type Trailing Edges

Viscous Damping of Turbulent Flow Pressures in the Brush-Region 3 10.00 3 4.00 3 10.00 5 11.71 35.7 4 27.2 1.023 2,3 0 36.5

45 Noise Concept #3.17 Trailing Edge Serrations

Reduced Acoustic Scattering, Reduced SpanWise Correlation 3 10.00 3 4.00 3 10.00 5 11.71 35.7 4 27.2 1.023 2,3 0 36.5

31 Noise Concept #3.03 Chevrons

Scaloped Edges Applied Jet Exhaust Nozzles 2.22 3 4.00 3 10.00 7 17.50 33.7 9 10.0 1.080 1,3,4 9 36.4

98 Prop Concept #4.30 Combined Steam Cycle

Combined gas plus steam turbine takes advantage of waste heat to produce work, improving thermal efficiency. 3 10.00 1 1.57 5 19.52 3 6.00 37.1 2 46.1 0.961 4 0 35.6

63 Noise Concept #3.35 Linear Acoustic Liner 1 3.92 1 1.57 3 10.00 7 17.50 33.0 9 10.0 1.080 3,4 8 35.6

65 Noise Concept #3.37 Double DOF Liner 1 3.92 1 1.57 3 10.00 7 17.50 33.0 9 10.0 1.080 3,4 9 35.685 Prop Concept #4.17 4-Shaft Turbine

Increasing the ability of the turbomachinery to rotate at ideal speeds and stage matching increases efficiencies. 5 19.52 3 4.00 3 10.00 5 11.71 45.2 1 100.0 0.783 4 4 35.4

57 Noise Concept #3.29 Active Noise Cancellation

Destructive Interference Introduced by a Powered Source 3 10.00 3 4.00 3 10.00 5 11.71 35.7 2 46.1 0.961 3 0 34.3

62 Noise Concept #3.34

Landing Gear Assembly "Plasma" Fairings

Electrodes Which Introduce Energy into the Flow Field Surrounding the Gear Assembly 3 10.00 3 4.00 3 10.00 5 11.71 35.7 2 46.1 0.961 3 0 34.3

49 Noise Concept #3.21 Slat Cove Fillers Elimination of the Slat-Gap 3 10.00 1 1.57 3 10.00 5 11.71 33.3 4 27.2 1.023 2,3 -3 34.1

53 Noise Concept #3.25 Continuous Mold-Line Linkages Elimination of the Slat-Gap 3 10.00 1 1.57 3 10.00 5 11.71 33.3 3 34.2 1.000 2,3 15 33.347 Noise Concept #3.19 Flap Tip Fences Reduction of Vortex Generation 3 10.00 3 4.00 3 10.00 3 6.00 30.0 8 12.3 1.072 2,3 3 32.232 Noise Concept #3.04 Ejectors

Enhanced Jet Mixing, Increased Turbulence Through Entrained Air 1 3.92 3 4.00 3 10.00 5 11.71 29.6 8 12.3 1.072 1,3,4 8 31.8

79 Prop Concept #4.11 Thrust Vectoring Open Rotor

Flow deflection or engine rotation may allow STOL operation. 3 10.00 5 7.81 3 10.00 1 2.35 30.2 5 22.2 1.040 2,3,4 20 31.4

52 Noise Concept #3.24 Slat Tip Fences Reduction of Vortex Generation 3 10.00 3 4.00 3 10.00 3 6.00 30.0 5 22.2 1.040 2,3 3 31.2

94 Prop Concept #4.26 Cryogenic Motors

Motors cooled to superconducting temperatures would operate at extreme efficiencies. 3 10.00 3 4.00 5 19.52 3 6.00 39.5 1 100.0 0.783 1,2,4 28 30.9

37 Noise Concept #3.09 Porous Plug Designs Reduced Vortex Shedding 3 10.00 3 4.00 3 10.00 3 6.00 30.0 4 27.2 1.023 1,3,4 0 30.7

46 Noise Concept #3.18 Porous Flap Side Edges Reduced Vortex Shedding 3 10.00 3 4.00 3 10.00 3 6.00 30.0 4 27.2 1.023 2,3 0 30.7

NASA/CR—2010-216798 157
Figure E-1. QFD Worksheet (Sheet 6 of 6) 

p p g g ,

48 Noise Concept #3.20 Slat Trailing Edge Serrations

Reduced Acoustic Scattering, Shift to High Frequencies 3 10.00 3 4.00 3 10.00 3 6.00 30.0 4 27.2 1.023 2,3 -3 30.7

51 Noise Concept #3.23 Porous Slat Pressure Surfaces Reduced Vortex Shedding 3 10.00 3 4.00 3 10.00 3 6.00 30.0 4 27.2 1.023 2,3 0 30.7

55 Noise Concept #3.27 Micro-tab Features Reduced Flap Deflections/Maintained Lift 3 10.00 3 4.00 3 10.00 3 6.00 30.0 4 27.2 1.023 2,3 0 30.7

2 Airframe Concept #1.02 Fiber Metal Laminates

Glass fiber reinforcement bonded between aluminum sheet 3 10.00 3 4.00 3 10.00 3 6.00 30.0 3 34.2 1.000 1 6 30.0

3 Airframe Concept #1.03

Metallic Digital Direct Manufacturing

Step wise construction of 3D monolithic metallic components from wire or powder 3 10.00 3 4.00 3 10.00 3 6.00 30.0 3 34.2 1.000 1 2 30.0

10 Airframe Concept #1.10

Open-Celled and Stamped Composite Structures

Composite-based Trussed Beaded & Sinewave Substructures 3 10.00 3 4.00 3 10.00 3 6.00 30.0 3 34.2 1.000 1 0 30.0

11 Airframe Concept #1.11

Post-Buckled Composite Structures

Composite skins and panels capable of carrying load beyond initial buckling 3 10.00 3 4.00 3 10.00 3 6.00 30.0 3 34.2 1.000 1 0 30.0

13 Airframe Concept #1.13

Direct Manufacturing / Laser Sintering of Polymeric Materials 

Layered Fabrication of powdered polymers that form net parts using laser sintering technology 3 10.00 3 4.00 3 10.00 3 6.00 30.0 3 34.2 1.000 1 0 30.0

14 Airframe Concept #1.14

Carbon Nanotube Electrical Cables 

Metallic CNT offers electrical conductivity as good as a copper with the density less than 1.3 g/cc compared to 8.92 g/cc for copper. 3 10.00 3 4.00 3 10.00 3 6.00 30.0 3 34.2 1.000 0 0 30.0

50 Noise Concept #3.22 Slat Gap Liners Acoustic Absorption of High Frequencies 3 10.00 3 4.00 3 10.00 3 6.00 30.0 3 34.2 1.000 2,3 0 30.0

54 Noise Concept #3.26 Streamlined Slat Tracks Reduced Turbulence 3 10.00 3 4.00 3 10.00 3 6.00 30.0 3 34.2 1.000 2,3 3 30.078 Prop Concept #4.10 Thrust Vectoring Nozzles

Deflecting part or all of the exhaust allows STOL operation. 1 3.92 7 11.66 3 10.00 1 2.35 27.9 6 18.3 1.053 3,4 31 29.4

67 Noise Concept #3.39 Metal Foam Liner 1 3.92 1 1.57 3 10.00 5 11.71 27.2 6 18.3 1.053 3,4 8 28.6

NASA/CR—2010-216798 158
References 

NASA/CR—2010-216798 159
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106 Ko, A., Leifsson, L., Schetz, J., Mason, W., Grossman, B., Haftka, R., “MDO of a BlendedWing-Body Transport Aircraft with Distributed Propulsion,” 3rd AIAA Aviation Technology, 

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NASA/CR—2010-216798 170
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4. TITLE AND SUBTITLE

NASA N+3 Subsonic Fixed Wing Silent Efficient Low-Emissions Commercial Transport 

(SELECT) Vehicle Study 

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NNC08CA86C 

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6. AUTHOR(S)

Bruner, Sam; Baber, Scott; Harris, Chris; Caldwell, Nicholas; Keding, Peter; Rahrig, Kyle; 

Pho, Luck; Wlezian, Richard 

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Northrop Grumman Corporation 

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14. ABSTRACT

A conceptual commercial passenger transport study was performed to define a single vehicle for entry into service in the 2030 to 2035 

timeframe, meeting customer demands as well as NASA goals for improved fuel economy, NOx emissions, noise, and operability into

smaller airports. A study of future market and operational scenarios was used to guide the design of an advanced tube-and-wing 

configuration that utilized advanced material and structural concepts, an advanced three-shaft high-bypass turbofan engine, natural laminar 

flow technology, and a suite of other advanced technologies. This configuration was found to meet the goals for NOx emissions, noise, and 

field length. A 64 percent improvement in fuel economy compared to a current state-of-the-art airliner was achieved, which fell slightly 

short of the desired 70 percent goal. Technology maturation plans for the technologies used in the design were developed to help guide 

future research and development activities. 

15. SUBJECT TERMS

Passenger aircraft; Subsonic aircraft; Turbofan aircraft 

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