Engine Conceptual Design Studies

for a Hybrid Wing Body Aircraft

NASA/TM—2009-215680

November 2009

Michael T. Tong, Scott M. Jones, and William J. Haller

Glenn Research Center, Cleveland, Ohio

Robert F. Handschuh

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

ARL–TR–4719

GT2009–59568

U.S. ARMY

RESEARCH LABORATORY
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Engine Conceptual Design Studies

for a Hybrid Wing Body Aircraft

NASA/TM—2009-215680

November 2009

Michael T. Tong, Scott M. Jones, and William J. Haller

Glenn Research Center, Cleveland, Ohio

Robert F. Handschuh

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

ARL–TR–4719

GT2009–59568

U.S. ARMY

RESEARCH LABORATORY

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Turbo Expo 2009

sponsored by the American Society of Mechanical Engineers

Orlando, Florida, June 8–12, 2009
Acknowledgments

The authors would like to acknowledge Dr. Timothy Krantz of the U.S. Army Research Laboratory for the development of the 

gearbox and lubrication system weight correlation used in this study, and Mr. Christopher Snyder for his comments on the paper.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5285 Port Royal Road

Springfi eld, VA 22161

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 technical management. 
NASA/TM—2009-215680 1 

Engine Conceptual Design Studies for a Hybrid Wing Body Aircraft 

Michael T. Tong, Scott M. Jones, and William J. Haller 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Robert F. Handschuh 

U.S. Army Research Laboratory 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Worldwide concerns of air quality and climate change have 

made environmental protection one of the most critical issues 

in aviation today. NASA’s current Fundamental Aeronautics 

Research program is directed at three generations of aircraft in 

the near, mid and far term, with initial operating capability 

around 2015, 2020, and 2030, respectively. Each generation 

has associated goals for fuel burn, NOx, noise, and field-length 

reductions relative to today’s aircrafts. The research for the 

2020 generation is directed at enabling a hybrid wing body 

(HWB) aircraft to meet NASA’s aggressive technology goals. 

This paper presents the conceptual cycle and mechanical 

designs of the two engine concepts, podded and embedded 

systems, which were proposed for a HWB cargo freighter. 

They are expected to offer significant benefits in noise 

reductions without compromising the fuel burn. 

Introduction 

More passengers and cargo are moved by air today than 

ever before, because of the global economy and worldwide 

connectivity. Over the next 15 to 20 years, the volume of air 

traffic is expected to at least double (for passenger traffic) or 

even triple (for cargo traffic) (Refs. 1 and 2). This robust 

growth rate causes growing concerns about the contribution 

that aircraft emissions will have on local air quality and global 

climate change. Chemical emissions of concern consist of 

anything that affects local air quality, global climate, or 

atmospheric ozone, including CO2, NOX, sulfur oxides, water 

vapor and particulates (Ref. 3). For carbon based fuels, there is 

a 1:1 relationship between the amount of fuel burned and the 

amount of CO2 generated. Aviation noise can have adverse 

impacts on property values, airport expansion, and prompts 

operational restrictions on existing runways that increase 

congestion, leading to travel and shipping delays (Ref. 4). It is 

generally recognized that significant improvement to the 

environmental acceptability of aircraft will be needed to 

sustain long term growth. The ability of the nation to benefit 

from continued growth in aviation depends on the 

development of future aircrafts that can meet demanding 

environmental and performance challenges. 

To achieve environmental protection that allows sustained 

long-term aviation growth, NASA has been engaged in the 

development of revolutionary aero-propulsion technologies 

and aircraft concepts with specific objectives to reduce aircraft 

fuel burn, noise, and NOx emissions while satisfying the field 

length constraints. Under the Subsonic Fixed Wing (SFW) 

project of its Fundamental Aeronautics Program, NASA’s 

aeronautics research is directed at three generations of aircraft 

in the near, mid and far term, with initial operating capability 

(IOC) around 2015, 2020, and 2030, respectively. Each 

generation has associated goals for reductions in noise, 

emissions, fuel burn, and field length relative to today’s 

aircraft. The three generations of aircraft are designated as 

‘N+1’, ‘N+2’, and ‘N+3’, respectively. The research for ‘N+2’ 

and ‘N+3’ are directed at enabling new vehicle configurations 

to meet NASA’s aggressive system-level goals. The ‘N+1’ 

and ‘N+2’ goals, as defined in the 2007 NASA Research 

Announcement request for proposal, are shown in Table 1. 

NASA funded a 1-year Phase-1 effort to study the potential 

of a Hybrid Wing Body type aircraft to meet the N+2 systemlevel goals. This study was to focus on the noise goal of –42 

dB relative to Federal Aviation Regulations Part 36 

(FAR 36) Stage 4 (–52 dB relative to Stage 3) while meeting 

the fuel goal of –25 percent relative to the current state-of-theart aircraft. Boeing Phantom Works, teamed with 

Massachusetts Institute of Technology (MIT) and University 

of California Irvine, proposed to perform the study on a 

freighter aircraft. Both Boeing and Airbus forecasted the 

demand for cargo air traffic will grow at a higher rate than 

passenger airliners in the next 20 years (Refs. 1 and 2). The 

team was chosen to conduct the study. 

Boeing, with its extensive background in blended wing 

body type aircraft, proposed two engine concepts for a hybrid 

wing body (HWB) freighter aircraft, for the 2020 timeframe—

the conventional pylon-mounted ‘podded’ and the futuristic 

‘embedded’ systems. The HWB configurations with podded 

and embedded engines were designated as ‘N2A’ and ‘N2B’, 

respectively. 

The N2A podded engine configuration was considered to be 

‘lower risk’ for the 2020 timeframe, because of its low engine 

operability risk. The N2B with embedded engines was 

considered to be a ‘higher risk’ configuration, because of its 
NASA/TM—2009-215680 2 

TABLE 1.—NASA SUBSONIC FIXED WING SYSTEM-LEVEL GOALS 

 N+1 generation 

conventional 

IOC 2015 

N+2 generation 

hybrid wing 

IOC 2020 

Noise 

(cumulative below Stage 4) –32 dB –42 dB 

Landing-and-takeoff NOx

emissions (below CAEP/6) –60% –75% 

Aircraft fuel burn 

(relative to a

B737/CFM56) 

b

–15% c–25%

a

N+2 baseline changed to B777/GE90 in 2008 

b

–33% with laminar flow control 

c

–40% with laminar flow control 

N2A N2B 

(with podded engines) (with embedded engines) 

Figure 1.—HWB aircraft-engine configurations. 

complexities associated with closely coupled engine/airframe 

and boundary layer ingestion inlets. The closely coupled 

engine/airframe has the potential to reduce the engine-airframe 

integration penalties. The N2B was to be derived from the 

“Silent Aircraft” (Refs. 5 and 6). The HWB aircraft-engine 

configurations with two types of engine are shown in Figure 1. 

Under the contractual agreement, NASA Glenn Research 

Center (GRC) agreed to perform engine conceptual design 

studies and provide the engine data to support Boeing’s effort. 

The design studies were for four podded engines with fan 

pressure ratios (FPR) of 1.4, 1.5, 1.6, and 1.7, and one 

embedded engine with FPR1.5 (mutually agreed to be the 

same as the Silent Aircraft engine). This paper presents the 

conceptual cycle and mechanical designs of the two engine 

concepts proposed for the ‘N+2’ generation freighter aircraft. 

Hybrid Wing Body (HWB) Aircraft 

A hybrid wing body aircraft is an alternative airframe 

design in which the fuselage blends seamlessly with the wings 

to form a hybrid flying-wing configuration (Ref. 7). It also 

incorporates many design features from the conventional ‘tube 

with wings’ aircraft. Because of its high-lift wings and wide 

airfoil-shaped body (thus better aerodynamic efficiency), the 

HWB aircraft reduces the drag and fuel burn. Fully integrating 

the HWB airframe and the engines, e.g., embedded engines, 

will allow the aerodynamic efficiency to be maximized, which 

would further improve the aircraft performance. Also, if the 

engines are installed above the wing, the engine noise will be 

shielded by the aircraft’s wide body and wing span and thus 

the aircraft will potentially operate quieter than the 

conventional aircraft. 

A very quiet HWB airplane would not be limited by current 

operational curfews, such as night operations into noisesensitive airports. The flexibility of operations, in combination 

with the worldwide trend towards widespread use of just-intime delivery, would further stimulate the cargo growth and 

the demand for freighter aircrafts. 

Aircraft Mission Requirements 

Boeing defined the mission requirements for a HWB cargo 

freighter aircraft. They are: 

 Payload of 103,000 lb; 

 Range of 6000 nm; 

 35000 ft initial cruise altitude or higher; 

 Time to climb through 31,000 ft not greater than 

30 min; 

 Cruise Mach number of 0.8; 

 Field length of 10,000 ft or less 

Propulsion System Design 

Propulsion System Design Requirements 

Based on the mission requirements, the propulsion system 

requirements were defined as follows: 

For the podded twinjet engine system: 

 Aerodynamic design point (ADP): Mach number 0.8 at 

31,000 ft; 

 Thrust (per engine) = 15000 lb at International standard 

atmosphere (ISA +0) 

 Rolling takeoff (RTO) at Mach no. 0.25, sea level: 

thrust (per engine) = 54000 lb (at ISA+15C/ISA+27F); 

for the embedded system (three engines, 9 fans): 

 Aerodynamic design point (ADP): Mach number 0.8 at 

31,000 ft; thrust (per engine) = 10000 lb at ISA+0) 

 Rolling takeoff at Mach no. 0.25, sea level: thrust (per 

engine) = 36000 lb (at ISA+15C/ISA+27F); 

Engine Cycle Design 

Cycle design involves simultaneously solving aerodynamic 

design point and off-design parameters. Four podded engines 

with FPR of 1.4, 1.5, 1.6, and 1.7 were modeled. Of these 

engines, the FPR1.4 and FPR1.5 engines were geared; the 

others were direct-drive. One embedded engine with FPR of 

1.5 was modeled. It has three propulsion modules that were 

each composed of a gas generator that drove an inline fan and 

two additional outboard fans through a mechanical drive train. 

The NASA software tool, NPSS (Numerical Propulsion 

System Simulation) (Refs. 8 and 9), was used for this task that 

ultimately calculated engine thrust and specific fuel 

consumption for each of the engines. All engines were 

developed with the same ADP (Mach number, altitude, thrust). 
NASA/TM—2009-215680 3 

The ADP was selected to represent a nominal top-of-climb 

(TOC) condition for the hybrid wing airframe cargo freighter. 

Inlet mass flow for each engine was selected to achieve the net 

thrust requirement at ADP and bypass ratio was set to achieve 

an extraction ratio (ratio of total pressures for bypass nozzle

and core nozzle) of 1.25 at the ADP for all engines. In 

addition to meeting a thrust target at TOC conditions, a sealevel rolling takeoff thrust target was also met by adjusting 

design point burner fuel-to-air ratio. 

A maximum high-pressure turbine (HPT) inlet temperature 

of 3460 °R and maximum HPT rotor inlet temperature of 

3310 °R (with cooling air) were assumed, reflecting the use of 

advanced high temperature materials. Also, a maximum lowpressure turbine (LPT) rotor inlet temperature of 2460 °R was 

used to eliminate the LPT cooling. 

Assumptions for fan, low pressure compressor (LPC), 

efficiencies were based on technology trend curves recently 

developed by the Aerospace Systems Design Lab (ASDL) at 

Georgia Tech for use in the FAA’s Environmental Design 

Space (EDS) system (Ref. 10). These curves have been 

reviewed by the EDS Independent Review Group, which 

includes industry representatives and is shown in Figures 2 

and 3. For the high pressure compressor (HPC), a constant 

polytropic efficiency of 91.5 percent was assumed for all the 

engines. For the FPR1.4 and FRP1.5 engines, a variable area 

fan nozzle was needed to achieve the targeted 20 percent surge 

margin across the operating envelope. For the FPR1.6 and 

FPR1.7 engines, an acceptable surge margin was achievable 

with fixed geometry nozzles and the extra weight of a variable 

area nozzle was not justified. 

General cycle characteristics of the podded engines are 

shown in Table 2. For the embedded engine, they are shown in 

Table 3. These data were generated with the inlet pressure 

recoveries provided by Boeing. For the podded engines, the 

Figure 2.—Variation of fan efficiency with pressure ratio. 

Figure 3.—Variation of LPC efficiency with pressure ratio. 

inlet pressure recoveries were 0.998 at the ADP. For the 

embedded engine, they were 0.946 and 0.960 (with boundary 

layer ingestion) for the center and the side inlets, respectively 

Engine Mechanical Design. 

TABLE 2.—GENERAL CYCLE CHARACTERISTICS OF PODDED ENGINE MODELS 

FPR = 1.40 FPR = 1.50 FPR = 1.60 FPR = 1.70 FPR = 1.40 FPR = 1.50 FPR = 1.60 FPR = 1.70 

 SLS 

(ISA+27 °F) 

ADP 

(ISA+0) 

SLS 

(ISA+27 °F) 

ADP 

(ISA+0) 

SLS 

(ISA+27 °F) 

ADP 

(ISA+0) 

SLS 

(ISA+27 °F) 

ADP 

(ISA+0) 

Fan Pressure 

Ratio (FPR) 

1.35 1.40 1.46 1.50 1.57 1.60 1.70 1.70 

Bypass Ratio 

(BPR) 

17.41 16.55 12.86 12.41 9.94 9.76 7.91 7.93 

Overall Pressure 

Ratio (OPR) 

43.7 48.4 43.6 46.4 43.5 44.9 43.6 43.6 

Net Thrust per 

engine, lb 

74859 15001 71838 15001 69755 15001 68256 15001 

TSFC, lb/(lb-h) 0.220 0.474 0.253 0.495 0.283 0.516 0.313 0.537 

HPT inlet temp. 

(T4), R 

3460 3048 3460 3000 3460 2969 3460 2947 

HPT rotor inlet 

temp. (T41), R 

3310 2913 3310 2868 3310 2838 3310 2817 

LPT rotor inlet 

temp., R 

2460 2144 2460 2109 2460 2084 2460 2067 

SLS = Sea level static 

ADP = Aerodynamic design point 

ISA = International standard atmosphere 
NASA/TM—2009-215680 4 

TABLE 3.—EMBEDDED ENGINE CYCLE PARAMETERS 

 SLS 

(ISA+27 °F) 

ADP 

(ISA+0) 

Fan Pressure Ratio (FPR) 1.49 1.50 

BPR (Center or core engine only) 3.2 3.1 

Effective BPR (Core and outboard engines) 11.5 11.3 

Overall Pressure Ratio (OPR) 45 46 

Net thrust per engine (3 fans), lb 49060 10000 

TSFC, lb/(lb-h) 0.288 0.564 

HPT inlet temp. (T4), R 3460 3010 

HPT rotor inlet temperature (T41), R 3310 2876 

LPT rotor inlet temperature, R 2460 2113 

The podded-engine system for the HWB cargo freighter is a 

twinjet (2 engines) system. For this system, four engine 

designs with fan pressure ratios (FPR) of 1.4, 1.5, 1.6, and 1.7 

were modeled. The basic common engine architecture for 

these engines is a two spool turbofan. Of these engines, the 

FPR1.4 and FPR1.5 engines were geared to reduce the number 

of LPC and LPT stages; the others were direct-drive. 

The embedded-engine system for the HWB cargo freighter 

is a 3-engine configuration with a total of 9 fans. Each 

embedded engine is composed of a gas generator (core engine) 

that drove an inline fan and two additional outboard fans 

through a mechanical drive train. For this concept, one engine 

design with FPR of 1.5 was modeled. 

The NASA software tool WATE (Weight Analysis of Gas 

Turbine Engines) (Refs. 11, 12, and 13) was used to create 

engine architectures that could achieve the engine 

thermodynamic cycle detailed in the previous section. Since 

WATE’s original release in 1979, substantial improvements 

have been made to enhance its capability and improve its 

accuracy. Many of the empirical relationships have been 

replaced with analytical weight and dimension calculations. 

An approach is used where the stress level, maximum 

temperature and pressure, material, geometry, stage loading, 

hub-tip ratio, blade/vane counts, and shaft speed are used to 

determine the component weight. An updated gearbox-weight 

correlation is also included in the code. 

The cycle data required for WATE execution, such as 

airflow, temperatures, and pressures, pressure ratios, bypass 

ratios, etc., was derived from NPSS output. Both the ADP and 

off-design cases were used to encompass the maximum 

performance level required for each engine component. This 

data, the material properties, and design rules for geometric, 

stress, and turbomachinery stage-loading limits were used to 

determine the acceptable engine layout. 

Advanced materials were assumed to accommodate higher 

engine operating temperatures and to reduce the weight. A 

complete summary of the advanced engine materials assumed 

is shown in Table 4. 

Both highly-loaded and conventional turbomachinery stageloading cases were studied. Using highly-loaded 

turbomachinery can reduce the number of compressor and 

turbine stages, reducing component and engine weights and 

lengths, but with a trade-off of component and overall 

efficiency. Based on the results, it was mutually agreed 

between GRC and Boeing that the small differences in weight 

(<2 percent) and overall dimensions (<6 percent in length) 

would not compensate for the overall efficiency degradation 

with the highly-loaded turbomachinery. In this paper, only the 

results based on conventional turbomachinery loadings are 

presented. 

For the podded engines the core nozzles were 

axisymmetric, and variable area fan nozzles were used for the 

FPR1.4 and FPR1.5 engines. It was assumed that the variablearea geometry (to be actuated by shape memory alloy) would 

increase the nozzle weight by 10 percent (Ref. 14). For the 

embedded engines, vectoring 2–D variable-area nozzles were 

used. The length of the nozzle was set at 2 fan diameters to 

provide the space for the acoustic liners. Tables 5 and 6 

summarize the parameters of the podded and embedded 

engines. The engine layouts for estimating performance and 

weights are shown in Figures 4 and 5. For the embedded 

engine, the boundary-layer-ingestion inlet and nacelles were 

considered part of the airframe and were designed by Boeing. 

TABLE 4.—ADVANCED ENGINE MATERIAL ASSUMPTIONS AND THEIR APPLICATIONS 

Component Blade Vane Disk Case 

Fan Polymer matrix 

composite 

Polymer matrix 

composite 

Current state-of-the-art 

materials 

Polymer matrix composite wrapped 

by Zylon 

LPC Titanium aluminide Titanium aluminide Current state-of-the-art 

materials 

Polymer matrix composite 

HPC 

(Hot section) 

Titanium aluminide Titanium aluminide Current state-of-the-art 

materials 

Titanium metal matrix composite 

HPT and LPT 5th generation 

nickel-based alloy 

5th generation 

nickel-based alloy 

Nickel-based powder 

metallurgy alloy 

Current state-of-the-art materials 

Inlet/Nacelle N/A N/A N/A Polymer matrix composite 
NASA/TM—2009-215680 5 

TABLE 5.—PRINCIPAL MECHANICAL PARAMETERS FOR THE PODDED ENGINES 

 FPR1.4 FPR1.5 FPR1.6 FPR1.7 

Configuration Two-spool geared 

turbofan 

Two-spool geared 

turbofan 

Two-spool direct drive 

turbofan 

Two-spool direct drive 

turbofan 

Fan dia., in. 126.6 115.1 106.8 100.3 

Fan blade/vane counts 18/46 18/46 18/46 18/46 

Max. fan tip speed, ft/sec 1119 1297 1450 1580 

Fan hub/tip ratio 0.31 0.31 0.31 0.31 

Fan stage loading* 0.28 0.25 0.24 0.23 

LPC stages 2 2 5 4 

HPC stages 9 9 8 8 

HPC min. blade ht., in. 0.62 0.69 0.73 0.79 

HPT stages 2 2 2 2 

LPT stages 3 3 6 5 

Fan nozzle type Variable area Variable area Fixed area Fixed area 

Total engine pod wt., lb 19007 16191 15513 13314 

Bare engine length, in. 178.4 166.3 185.6 164.8 

TABLE 6.—PRINCIPAL MECHANICAL PARAMETERS 

FOR THE EMBEDDED ENGINE 

Configuration Multiple fan system 

Fan dia., in. .......................................................................................... 56 

Fan blade/vane counts .................................................................... 18/44 

Max. fan tip speed, ft/sec ................................................................ 1297 

Fan hub/tip ratio ............................................................................... 0.31 

Fan stage loadinga

 ............................................................................ 0.25 

LPC stages ............................................................................................. 5 

LPC blade/vane counts ............................................................... 193/360 

HPC stages ............................................................................................ 9 

HPC blade/vane counts .............................................................. 554/824 

HPC min. blade ht. (in.) ................................................................... 0.68 

HPT stages ............................................................................................ 2 

HPT blade/vane counts ................................................................ 111/67 

LPT stages ............................................................................................. 5 

LPT blade/vane counts ............................................................... 657/436 

Nozzle type ................................................................. 2–D variable area 

Engine weight (includes accessories, 

with no transmission), lb .............................................................. 12,652 

Transmission and lubrication system weight, lb ........................... 1,139 

Total engine weight (excludes inlet), lb ....................................... 13791 

a

Fan stage loading =

2 2 (1 / ) 2

1U h t

h

t 



∆h = change in stagnation enthalpy 

Ut = blade tip speed 

h/t = blade hub-to-tip ratio 
NASA/TM—2009-215680 6 

Figure 4.—FPR1.6 podded engine internal layout (dimensions in inches). 

Figure 5.—Embedded engine internal layout. 
NASA/TM—2009-215680 7 

Figure 6.—The gear-drive system configuration. 

Transmission Design for the Embedded Engine 

Each embedded engine is composed of a gas generator that 

drove an inline fan and two additional fans through a 

mechanical drive train. The mechanical drive train was 

designed to be powered from the low pressure turbine (LPT) 

through angle gearboxes to adjacent fans. The gear-drive 

system configuration is shown in Figure 6. 

The gearboxes were designed using the calculation 

procedure for spiral bevel gears via the American Gear 

Manufacturers Association (AGMA, (Ref. 15)). The load 

(power) was assumed to be split equally between the three 

fans. Therefore the gearbox driven directly by the power 

turbine was designed to transmit ~35 khp (Figure 6, Gearbox 

#1), or 2/3 of the power, and then split the power to the 

adjacent gearboxes to drive the two parallel outboard fans 

(Figure 6, Gearboxes #2). The gearbox arrangements also 

considered overall size to minimize the cross-sectional area 

down stream of the turbine and fans. Gearbox #1 was 

penalized during the design process since the pinion drives 

two gears. State-of-the-art materials and manufacturing 

processes would be required for all gearbox system 

components. The gearing design parameters are shown in 

Table 7. 

Figure 7.—Rotorcraft transmission and 

lubrication system weight data. 

An empirical correlation, shown in Figure 7, was used to 

calculate the weight of the transmission and lubrication 

system. The correlation was developed based on actual weight 

data from over fifty rotorcrafts, tiltrotors, and turboprop 

aircrafts. They are also plotted in Figure 7. Using this 

parametric fit permitted gearbox weight to be estimated for the 

purposes of this study. 

Aircraft Mission and Sizing Studies 

With the engine data provided by GRC, Boeing used its 

BIVDS (Boeing Integrated Vehicle and Design System) tool 

suite to perform airplane mission and sizing analyses, based 

on an 11112-km (6000 n mi) economic mission. The results 

have been reported in Reference 16. For the podded engines, 

they are summarized in Table 8. 

TABLE 8.—MISSION AND SIZING RESULT COMPARISONS 

FOR THE PODDED ENGINES 

Ground rules: 

 6000 nm range 

 30 min time to climb through 31000 ft 

 35,000 initial cruising altitude (ICA) 

 Hot-day takeoff: ISA+27 °F 

Fan pressure 

ratio 

1.4 1.5 1.6 1.7 

Maximum 

takeoff gross 

weight, lb 

464,700 460,700 461,500 463,700 

Payload, lb 103,000 103,000 103,000 103,000 

Static sea level 

thrust, lb 

74,862 71,837 69,757 68,258 

Fuel burn, lb 118,573 120,939 125,051 129,127 

Engine outfield 

length, ft 

6,214 5,942 6,196 6,320 

TABLE 7.—THE GEARING DESIGN PARAMETERS 

Number of teeth pinion/gear 48/25 27/52 

Diametral pitch, 1 in. 2.5 3.0 

Spiral angle, degrees 25 25 

Face width, in. 3.5 3.25 

Outside diameter, in. 19.41/10.94 9.79/17.51 

Pinion bending stress, ksi 64.3 67.2 

Gear bending stress, ksi 67.4 68.6 

Contact stress, ksi 184.2 182.2 

Pitch line velocity, ft/min. 25178 22660 
NASA/TM—2009-215680 8 

Boeing used the fuel-burn trend band for current cargo 

freighters (B767-300ER, A330-200, etc.) for the comparison. 

It showed that the N2A (with FPR1.6 podded engines) 

exceeds N+2 fuel burned goal at –29 percent. Although both 

the FPR1.4 and FPR1.5 geared engines had lower fuel burn, 

the FPR1.6 engine was deemed to have lower risk for the 2020 

IOC time frame. It was selected for the noise study. With the 

embedded engines, the N2B met the fuel-burn goal at 

–25 percent. Those results are summarized in Table 9. 

TABLE 9.—MISSION AND SIZING RESULTS 

FOR THE EMBEDDED ENGINE 

Ground rules: 

 6000 nm range 

 30 min time to climb through 31000 ft 

 35,000 ICA 

 Hot-day takeoff: ISA+27 °F 

Fan pressure ratio ................................................................................... 1.5 

Maximum takeoff gross weight, lb ................................................ 477,400 

Payload, lb ...................................................................................... 103,000 

Static sea level thrust, lb .................................................................. 48,320 

Fuel burn, lb ................................................................................... 130,300 

Engine outfield length, ft .................................................................... 5436 

Aircraft and Engine Noise Studies 

Subsequent noise studies were also conducted by Boeing 

and MIT. The methodology and results are reported in 

Reference 16. The noise estimate for the N2A was shown to 

be –47 dB below Stage 3 (or –37 dB below Stage 4), within 

5 dB of the N+2 goal. For the N2B, the noise was shown to be 

–26 dB below Stage 3 (or –16 dB below Stage 4). Based on 

the results, Boeing concluded that the N+2 noise goal is 

achievable with N2A configuration, with increased jet 

shielding, increased climb speed, additional focus on landing 

gear fairings, and with continuing R&D on HWB type aircraft. 

For the N2B, increasing the duct treatment (e.g., with acoustic 

tiles) and reducing the jet velocity will help it move towards 

the N+2 noise goal. A part of the continuing R&D is the need 

to further improve noise prediction methodologies, especially 

for an embedded engine. 

Summary 

NASA GRC conducted engine conceptual design studies on 

two engine concepts, podded and embedded systems, that 

were proposed for a HWB freighter aircraft for the ‘N+2’ 

timeframe. The results were provided to Boeing Phantom 

Works to support its investigation to develop a HWB subsonic 

freighter configuration with noise prediction methods to meet 

the NASA Subsonic Fixed Wing N+2 noise and fuel burn 

reduction goals. Based on its Phase 1 results, Boeing has 

concluded that the N+2 fuel burn and noise goals are 

achievable on a hybrid wing type vehicle, with continuing 

R&D on HWB type aircraft and improvement of noise 

prediction methodologies. 

References 

1. The Boeing Company, “Current Market Outlook 2008-

2027,” July 9, 2008. Retrieved on November 10, 2008 

from: http://www.boeing.com/commercial/cmo/pdf/ 

Boeing_Current_Market_Outlook_2008_to_2027.pdf 

2. Airbus, “Global Market Forecast 2007–2026,” 2007. 

Retrieved on November 10, 2008 from: 

http://www.airbus.com/fileadmin/documents/gmf/PDF_dl/

00-all-gmf_2007.pdf 

3. National Research Council, “Securing the Future of U.S. 

Air Transportation: A System in Peril,” Washington, DC, 

National Academy Press, 2003. 

4. National Research Council, “For Greener Skies,” National 

Academy Press, 2002. 

5. Cambridge-MIT Institute, “Silent Aircraft Conceptual 

Design,” November 6, 2006. 

6. Hall, C.A. and Crichton D., “Engine Design Studies for a 

Silent Aircraft,” Journal of Turbomachinery, 129, July, 

2007. 

7. Liebeck, R.H., “Design of the Blended Wing Body 

Subsonic Transport,” Journal of Aircraft, 41, No. 1, 

January—February, 2004. 

8. NASA-Industry Cooperative Effort, “Numerical Propulsion 

System Simulation User Guide and Reference,” Software 

Release NPSS 1.5.0, May 7, 2002. 

9. Lytle, J.K., “The Numerical Propulsion Simulation: An 

Overview,” NASA/TM—2000-209915. 

10. Kirby, M.R., “Environmental Design Space, Tool 

Developments Enhancements,” Semi-annual progress 

report, US Department of Transportation/Federal Aviation 

Administration, Project No. 1606A75, Contract No. 07-CNE-GIT-1, May 2008. 

11. Onat, E. and Klees, G.W., “A Method to Estimate Weight 

and Dimensions of Large and Small Gas Turbine Engines,” 

NASA CR–159481, 1979. 

12. Tong, M.T., Halliwell, I., Ghosn, L.J., “A Computer Code 

for Gas Turbine Engine Weight and Life Estimation,” 

ASME Journal of Engineering for Gas Turbine and Power, 

volume 126, no. 2, April 2004. 

13. Tong, M.T., Naylor, B.A., “An Object-Oriented Computer 

Code for Aircraft Engine Weight Estimation,” GT2008–

50062, ASME Turbo-Expo 2008, June 9–13, 2008. 

14. Geiselhart, K., Berton, J.J., Tong, M.T., “Shape Memory 

Alloy Variable Area Fan Nozzle,” NASA REVCON ISAT 

Report, April 2001. 

15. American Gear Manufacturers Association Standard, 2003-

A86,“Rating the Pitting Resistance and Bending Strength 

of Generated Straight Bevel, Zerol Bevel, and Spiral Bevel 

Gear Teeth,” January 1992. 

16. Kawai, R. and Brown, D., “Acoustic Prediction 

Methodology and Test Validation for an Efficient LowNoise Hybrid Wing Body Subsonic Transport,” Phase I 

Final Report PWDM08-006A, NASA Contract Number 

NNL07AA54C, October, 2008. 
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Engine Conceptual Design Studies for a Hybrid Wing Body Aircraft 

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14. ABSTRACT

Worldwide concerns of air quality and climate change have made environmental protection one of the most critical issues in aviation today. 

NASA’s current Fundamental Aeronautics Research program is directed at three generations of aircraft in the near, mid and far term, with 

initial operating capability around 2015, 2020, and 2030, respectively. Each generation has associated goals for fuel burn, NOx, noise, and 

field-length reductions relative to today’s aircrafts. The research for the 2020 generation is directed at enabling a hybrid wing body (HWB) 

aircraft to meet NASA’s aggressive technology goals. This paper presents the conceptual cycle and mechanical designs of the two engine 

concepts, podded and embedded systems, which were proposed for a HWB cargo freighter. They are expected to offer significant benefits in 

noise reductions without compromising the fuel burn. 

15. SUBJECT TERMS

Hybrid wing body; Fuel burn; Noise; Emissions 

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