January 2011 

NASA/TM–2011-216883 

Refined Exploration of Turbofan Design 

Options for an Advanced Single-Aisle Transport 

Mark D. Guynn 

Langley Research Center, Hampton, Virginia 

Jeffrey J. Berton, Kenneth L. Fisher, William J. Haller, and Michael T. Tong 

Glenn Research Center, Cleveland, Ohio 

Douglas R. Thurman 

Army Research Laboratory, Cleveland, Ohio 
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January 2011 

NASA/TM–2011-216883 

Refined Exploration of Turbofan Design 

Options for an Advanced Single-Aisle Transport 

Mark D. Guynn 

Langley Research Center, Hampton, Virginia 

Jeffrey J. Berton, Kenneth L. Fisher, William J. Haller, and Michael T. Tong 

Glenn Research Center, Cleveland, Ohio 

Douglas R. Thurman 

Army Research Laboratory, Cleveland, Ohio 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

Acknowledgments 

The authors would like to thank the Subsonic Fixed Wing Project of NASA’s 

Fundamental Aeronautics Program for supporting this study. 

The use of trademarks or names of manufacturers in this report is for accurate reporting and does not 

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the 

National Aeronautics and Space Administration. 
iii 

Contents 

List of Tables ...............................................................................................................................................iv

List of Figures............................................................................................................................................... v

List of Acronyms, Symbols, and Abbreviations..........................................................................................vi

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

1.0 Introduction.............................................................................................................................................1

2.0 Study Objectives and Approach..............................................................................................................2

3.0 Modeling and Analysis Methodology.....................................................................................................3

3.1 Propulsion System Modeling............................................................................................................................3

3.2 Aircraft Sizing Analysis ...................................................................................................................................4

3.3 Noise Analysis..................................................................................................................................................5

4.0 Study Engine Designs.............................................................................................................................6

4.1 Engine Trade Space..........................................................................................................................................6

4.2 Engine Comparison ..........................................................................................................................................9

5.0 Aircraft Sizing and Performance Results..............................................................................................12

6.0 Design Trade-Offs ................................................................................................................................23

7.0 Concluding Remarks.............................................................................................................................25

8.0 References ......................................................................................................... 26
iv 

List of Tables 

Table 1. Revised Engine Design Ground Rules and Assumptions ...............................................................................4

Table 2. Engine Trade Space........................................................................................................................................6

Table 3. General Characteristics of Low Work Engines ..............................................................................................8

Table 4. General Characteristics of High Work Engines..............................................................................................8

Table 5. Aircraft Sizing Results for Low Work Engines (162 Passenger, 3250 nm Design Mission) .......................14

Table 6. Aircraft Sizing Results for High Work Engines (162 Passenger, 3250 nm Design Mission).......................15

Table 7. Weight, Fuel, Emissions, and Noise Trade-Offs ..........................................................................................24
v 

List of Figures 

Figure 1. NASA Subsonic Fixed Wing Project's system level metrics and goals. .......................................................2

Figure 2. Relationship between bypass ratio and design fan pressure ratio, TOC conditions. .....................................9

Figure 3. Variation of TSFC with engine type and design fan pressure ratio, TOC conditions. ................................10

Figure 4. Variation of nacelle maximum diameter with engine type and design fan pressure ratio. ..........................10

Figure 5. Variation of engine+nacelle weight with engine type and design fan pressure ratio. .................................11

Figure 6. Variation of engine thrust-to-weight ratio with engine type and design fan pressure ratio.........................12

Figure 7. Variation of aircraft empty weight with engine type and design fan pressure ratio. ...................................16

Figure 8. Variation of aircraft ramp weight with engine type and design fan pressure ratio......................................17

Figure 9. Variation of block fuel with engine type and design fan pressure ratio. .....................................................18

Figure 10. Variation of total mission NOX emissions with engine type and design fan pressure ratio.......................19

Figure 11. Variation of LTO NOX emissions with engine type and design fan pressure ratio (Dp/Foo)......................19

Figure 12. LTO NOX margin relative to CAEP6 regulatory limit. .............................................................................20

Figure 13. Variation of NOX emissions per LTO cycle with engine type and design fan pressure ratio....................20

Figure 14. Potential block fuel reductions from application of ASAT technologies. .................................................21

Figure 15. Potential ramp weight reductions from application of ASAT technologies. .............................................21

Figure 16. Potential total NOX reductions from application of ASAT technologies. .................................................22

Figure 17. Potential LTO NOX reductions from application of ASAT technologies..................................................22

Figure 18. Potential noise reductions from application of ASAT technologies..........................................................23
vi 

List of Acronyms, Symbols, and Abbreviations 

ADP – Aerodynamic Design Point 

ANOPP – Aircraft Noise Prediction Program 

ASAT – Advanced Single-Aisle Transport 

BPR – Bypass Ratio 

CAEP – Committee on Aviation Environmental Protection 

EIS – Entry-Into-Service 

EPNL – Effective Perceived Noise Level 

FAR – Federal Aviation Regulations 

FLOPS – Flight Optimization System 

FPR – Fan Pressure Ratio 

HPC – High Pressure Compressor 

HPT – High Pressure Turbine 

ICAO – International Civil Aviation Organization 

ISA – International Standard Atmosphere 

LPC – Low Pressure Compressor 

LPT – Low Pressure Turbine 

LTO – Landing-Takeoff Cycle 

NPSS – Numerical Propulsion System Simulation 

OPR – Overall Pressure Ratio 

PDCYL – Point Design of Cylindrical-Bodied Aircraft

TOC – Top-Of-Climb 

TSFC – Thrust Specific Fuel Consumption 

UHB – Ultra-High Bypass Ratio 

WATE – Weight Analysis of Turbine Engines 
Abstract 

The desire for higher engine efficiency has resulted in the evolution of 

aircraft gas turbine engines from turbojets, to low bypass ratio, first 

generation turbofans, to today’s high bypass ratio turbofans. It is 

possible that future designs will continue this trend, leading to very-high 

or ultra-high bypass ratio engines. A comprehensive exploration of the 

turbofan engine design space for an advanced technology single-aisle 

transport (737/A320 class aircraft) was conducted previously by the 

authors and is documented in a prior report. Through the course of that 

study and in a subsequent evaluation of the approach and results, a 

number of enhancements to the engine design ground rules and 

assumptions were identified. A follow-on effort was initiated to 

investigate the impacts of these changes on the original study results. 

The fundamental conclusions of the prior study were found to still be 

valid with the revised engine designs. The most significant impact of the 

design changes was a reduction in the aircraft weight and block fuel 

penalties incurred with low fan pressure ratio, ultra-high bypass ratio 

designs. This enables lower noise levels to be pursued (through lower fan 

pressure ratio) with minor negative impacts on aircraft weight and fuel 

efficiency. Regardless of the engine design selected, the results of this 

study indicate the potential for the advanced aircraft to realize 

substantial improvements in fuel efficiency, emissions, and noise 

compared to the current vehicles in this size class. 

1.0 Introduction 

As aircraft manufacturers Boeing and Airbus continue to develop and mature new twin-aisle, wide 

body aircraft designs in the 210-350 seat class, for scheduled first deliveries in 2011 and 2013 

respectively, significant attention is also being paid to the potential for future new products in the smaller 

Boeing 737/Airbus A320 class. Options under consideration include taking advantage of evolving 

propulsion advances through a re-engining program in the fairly near term, to development of a 

completely new vehicle in the farther term. Airbus has chosen to initially pursue the near term benefits of 

re-engining, while Boeing continues to evaluate its options (ref. 1). Regardless of the path chosen, the 

fuel consumption and operating cost reductions necessary to make a new product economically viable 

will require substantial improvements in propulsion system efficiency. It is well known in aircraft 

propulsion system design that it is more efficient to generate thrust by accelerating a large mass of air a 

small amount than by accelerating a small mass of air a large amount; propulsive efficiency increases as 

the ratio of exhaust velocity to free stream velocity decreases. For a turbofan engine, this can be 

accomplished by reducing the fan pressure ratio (FPR), which decreases the amount of fan air stream 

acceleration, and increasing the fan mass flow (fan size) to maintain thrust. An increase in fan mass flow 

for a given core engine size leads to higher bypass ratio (BPR). The desire for higher engine efficiency 

has resulted in the evolution of aircraft gas turbine engines from turbojets (BPR=0), to low bypass ratio, 

first generation turbofans (BPR=1-2), to today’s high bypass ratio turbofans (BPR=5-10). It is possible 

that future designs will continue this trend, leading to very-high or ultra-high bypass ratio (UHB) engines. 

Reduced FPR has complementary benefits in lower engine noise due to the strong relationship between 

noise and the velocity of the air exiting the engine. Low pressure ratio fans also typically require lower tip 

speeds which can result in lower fan noise. Although there are fundamental noise and efficiency benefits 

to lowering FPR, there are typically weight and drag penalties which can potentially offset those benefits. 
2 

In addition, the larger fan diameter associated with lower FPR can lead to engine-airframe integration

issues. Only through analysis of the complete aircraft system can the best FPR for a given aircraft design 

be determined. 

Reference 2 describes a previous study conducted by the authors of this report in which the turbofan 

design trade space for an advanced single-aisle transport* (ASAT) was explored to investigate the impact 

of FPR and other engine design choices. The study was conducted to support the “N+1” goals of NASA’s 

Subsonic Fixed Wing Project (see Figure 1) by identifying the engine designs which provide the best 

opportunity to meet the goals for the N+1 timeframe. (“N” refers to the current generation of aircraft

flying today, “N+1” the next generation, “N+2” the second generation, and so forth.) The performance, 

noise, and emission characteristics of 48 different advanced engine/airframe combinations were evaluated 

and compared in reference 2. In developing the engines for that study, considerable attention was paid to 

ensuring consistent assumptions and ground rules were applied to all the engine models. However, when

the assumptions and ground rules were subjected to additional scrutiny through consultation with engine 

manufacturers, it was determined that because of the wide breadth of engine architectures and cycle 

parameters encompassed by the study, the “same ground rules across all engines” approach used 

previously was not always appropriate. A new set of modeling assumptions and ground rules was 

developed to better reflect the unique characteristics of the individual engines, and a subset of the original 

cases (12) was re-analyzed. This report presents the results of this new analysis and comparisons to the 

results of the original study documented in reference 2. 

Figure 1. NASA Subsonic Fixed Wing Project's system level metrics and goals. 

2.0 Study Objectives and Approach 

The primary objective of this study was to determine the impact of refinements to the engine design 

ground rules and assumptions used in reference 2 on the overall conclusions of that study. The basic 

analysis approach described in that report was retained; a series of analytical engine models were 

 

*

 “Single-Aisle Transport” is a common way to refer to a 737/A320 class airplane; although there are other types of 

single-aisle aircraft (e.g., regional jets). Even though it is possible that the future 737/A320 replacements designed 

by Boeing and Airbus will not have single-aisle passenger layouts, the term “single-aisle transport” will be used in 

this report to refer to a 737/A320 class airplane. 
3 

developed and applied to a common airframe model to assess the overall performance and noise 

characteristics. As in the prior study, the primary engine design parameter of interest was fan pressure 

ratio, which is directly related to bypass ratio. In addition, the impacts of fan drive approach (direct drive 

vs. geared) and compression work split (between the low pressure compressor (LPC) and high pressure 

compressor (HPC)) were investigated. Because the results of reference 2 showed that cruise Mach 

number and engine overall pressure ratio (OPR) did not significantly impact the relative trends with fan 

pressure ratio, those parameters were not varied in the current study. 

3.0 Modeling and Analysis Methodology 

3.1 Propulsion System Modeling 

The same propulsion system modeling tools were used for this study as in the prior work of reference 

2. Cycle analysis for the engines was performed with the NPSS (Numerical Propulsion System 

Simulation) code (refs. 3-5). Analysis of the aeromechanical characteristics and estimates of the engine 

weight (including fan gearbox if applicable) were performed with the WATE (Weight Analysis of 

Turbine Engines) code (refs. 6-8). Estimates for engine NOX emission indices (grams of NOX emitted 

from the engine per kilogram of fuel consumed by the engine) were obtained from a correlation 

developed by NASA combustor technologists during the latter stages of NASA’s Ultra-Efficient Engine 

Technology Program. Reference 2 provides more details on this empirical NOX correlation. 

All engines in the study are two-spool, separate flow, turbofans designed with the same Aerodynamic 

Design Point (ADP) and same OPR at the ADP. The ADP of 5,000 lb of thrust at Mach 0.8 and 35,000 ft 

was selected to represent a nominal top-of-climb (TOC) condition for the advanced technology airframe.

The ADP engine OPR of 42 is equivalent to the “Spiral 2” analysis in reference 2. Although the OPR is 

the same for all the engines, two different compressor work splits were considered. The “low work” 

engines have a lower pressure rise across the LPC (and a higher pressure rise across the HPC) compared

to the “high work” engines. Inlet mass flow for each engine was selected to achieve the net thrust 

requirement at ADP. The BPR was set by a jet velocity ratio (Vcore/Vbypass) of 1.6 at the design point. This 

differs from the prior work in which an extraction ratio of 1.25 (ratio of total pressures for bypass nozzle 

and core nozzle) was used to set BPR. In the prior study, design point burner fuel-to-air ratio was adjusted 

to achieve a sea-level static (SLS) target thrust of 23,000 lb (hot day, ISA+27°F) for all of the engines. 

Because of the wide variation in FPR and therefore thrust lapse rates, designing to equal SLS thrust did 

not provide equivalent takeoff performance for all the engines. For the current study, the engines were 

instead designed for equal thrust at a rolling takeoff condition (sea-level, Mach 0.25). The sea-level, 

Mach 0.25 thrust target was set to 17,500 lb based on the takeoff thrust required for a nominal 

configuration. Low FPR engine cycles generally require some type of variable geometry for proper 

operation across the flight envelope. Two approaches commonly considered are variable pitch fan blades

or a variable area fan exhaust nozzle. Because variable pitch fan blades present additional technological 

challenges, the use of a variable area nozzle was examined in this study. A variable area nozzle was 

applied when necessary to achieve the desired 20% fan surge margin throughout the operating envelope. 

The variable area nozzle was assessed an estimated weight penalty of 10% compared to an equivalent 

fixed-area design. 

In the prior analysis duct pressure loss assumptions were constant across all the engines. But because

of the large variation in geometry and flow conditions for the different engines, some variation in duct 

losses should be expected. In the current study the assumed bypass duct pressure loss was changed from a 

constant value to a function of FPR, with higher FPR engines having higher assumed bypass duct 

pressure losses due to the higher speed bypass duct flow. Another difference among the engine designs 
4 

requiring consideration was the different turbine geometries of the direct drive and geared engines. The 

turbines of the direct drive engines have larger radial variation, which should lead to higher inter-turbine 

duct pressure losses. This difference in inter-turbine duct losses was added in the revised engine 

modeling. The low pressure turbine (LPT) cooling philosophy is another area in which refinements were 

made. In the previous study, a maximum LPT rotor inlet temperature of 2460°R was used to allow the 

LPT to be uncooled with the use of advanced, high-temperature materials. However, in the high work 

engines, less work is done by the high-spool, leading to higher high pressure turbine (HPT) exit 

temperatures, and more cooling is required to meet a 2460°R LPT rotor inlet temperature limit. The 

constraint of an uncooled LPT was removed for the current study and a cooling analysis was performed 

for each engine to determine the amount of cooling air necessary to maintain acceptable HPT and LPT 

temperatures. The introduction of different cooling levels for the LPT also made the prior assumption of 

equal LPT adiabatic efficiency across all the engines invalid. For this study LPT efficiency was varied as 

a function of cooling level. A minor change to the engine designs was the assumption of a two stage HPT 

rather than the single stage assumed previously. The two stage design was felt to be more representative 

of likely industry designs (ref. 9), although this change had little impact on the engine weight and 

performance characteristics. 

Key differences between the engine design ground rules and assumptions of the current analysis and 

those in reference 2 are summarized in Table 1. 

Table 1. Revised Engine Design Ground Rules and Assumptions 

 Modeling in Reference 2 Current Study 

Engine Thrust Sizing ADP: 5,000 lb 

SLS: 23,000 lb 

ADP: 5,000 lb 

SL, M=0.25: 17,500 lb 

Bypass Ratio Extraction Ratio of 1.25 at ADP Jet Velocity Ratio of 1.6 at ADP 

Bypass Duct Losses Constant Function of FPR 

Inter-turbine Duct Losses Constant Different assumptions for direct 

drive and geared fan architectures 

Turbine Cooling Philosophy LPT temperature limited to allow 

uncooled LPT LPT cooling permitted 

LPT Adiabatic Efficiency Constant with constant LPT 

loading Function of LPT cooling level 

HPT Design Single stage Two stage 

3.2 Aircraft Sizing Analysis 

As in the prior study, the study engines were combined with an advanced technology, single-aisle 

commercial transport airframe model to determine overall system level performance. The advanced 

technology, “ASAT” airframe model is a derivative of a 737-800 like baseline model intended to be 

representative of a potential advanced technology replacement aircraft. The aircraft sizing and synthesis 

computer code FLOPS (Flight Optimization System) (ref. 10) was used as the primary aircraft level sizing 

and analysis tool. Special sizing considerations introduced by large diameter engines were addressed 

through simplifying assumptions and enhancements to the FLOPS analysis. Spreadsheet analyses were 

used to determine landing gear length, engine-out drag, and vertical tail size so that the impacts of large 

diameter engines could be properly captured. Enhancements to basic FLOPS capabilities were also made 
5 

in the structural weight and aerodynamics areas. The wing and fuselage structural weight estimates of 

FLOPS were replaced with estimates from PDCYL (ref. 11). PDCYL offers a less empirical, more 

analytical weight estimation methodology that is more sensitive to parameters such as engine weight and 

location. FLOPS aerodynamic predictions were enhanced through a model calibration process 

incorporating details of the 737-800 high speed and low speed aerodynamic performance. The primary 

airframe technology advancement assumed was extensive use of composite materials for the airframe 

structure. For the Boeing 787 currently in development, as much as 50% of the primary structure is made 

of composite materials (ref. 12). Other minor technology improvements based on the 787 included an 

increase in hydraulic system pressure and a slight drag reduction. Changes were also made to the design 

mission to reflect projected performance enhancements desired for an advanced aircraft in this vehicle

class. Cruise Mach was increased slightly to 0.8 (typical cruise Mach for the 737-800 is 0.785 (ref. 13)) 

and design range (with 32,400 lb payload) was increased from 3060 nm to 3250 nm. Other than a slight 

increase in wing sweep reflecting the higher cruise Mach number, the basic 737-800 geometry was not 

changed for the advanced technology, ASAT model. Further details of the airframe model development 

and calibration process are given in reference 2. 

3.3 Noise Analysis 

The same noise analysis methodology and tools described in reference 2 were used in the current 

effort. The primary tools used for the noise analysis included: NPSS for the engine cycle analysis; WATE 

for the engine aeromechanical and flowpath analysis; FLOPS for the aircraft trajectory simulation; and

ANOPP (Aircraft Noise Prediction Program) Level 26 (refs. 14, 15) for the source noise prediction and 

propagation. The NPSS and WATE codes were used to generate input data necessary for the ANOPP 

source noise modeling. Adjustments representing noise reduction technologies (discussed below) were 

made to the source noise spectra prior to propagation. ANOPP noise propagation modeling included 

spherical spreading, atmospheric attenuation, ground effects, reflections, and lateral attenuation. The 

Effective Perceived Noise Level (EPNL) was calculated at the noise certification points defined in FAR

Part 36 (ref. 16). EPNL is an integration of the ground observer perceived noise time history which 

depends on aircraft trajectory, noise spectra propagation, frequency integration, and tonal content and 

amplitude penalties. Validation of the noise analysis methodology using 737-800 certification data is 

described in reference 2. 

The same series of advanced noise reduction technologies were assumed as in the previous study. 

Chevrons were applied to all core nozzles and to all fixed-area bypass nozzles. Chevrons were not applied 

to variable area bypass nozzles due to potential conflict with the variable area nozzle design. Jet noise 

benefits of the nozzle chevrons were determined analytically using the 2004 Stone jet noise prediction

method in ANOPP (ref. 17). This method is based on 1997 acoustic measurements of chevron-equipped 

nozzles from NASA Glenn’s Aeroacoustic Propulsion Laboratory’s Nozzle Acoustic Test Rig freejet 

facility (ref. 18). Conventional inlet, interstage, and aft fan duct liners were applied to reduce fan inlet 

and discharge noise. The benefits of these liners were modeled by applying an acoustic suppression 

“map” of 1/3rd octave band sound pressure level decrements to the hardwall fan source spectra predicted 

by ANOPP. The liner suppression map was based on measured acoustic data of 22-inch diameter fan test 

articles in NASA Glenn’s 9×15 Low Speed Wind Tunnel (ref. 19). In addition to conventional liners, two 

advanced technologies were applied for fan noise reduction; soft vane stators and over-the-rotor foam 

metal treatment (refs. 20, 21). Both of these technologies are applications of acoustic treatment in areas of 

the engine which currently do not have treatment, the fan vanes and above the fan rotor tips. Acoustic

tests of both of these technologies were conducted at NASA Glenn in 2008. Airframe noise reduction 

technologies assumed included innovative slat cove designs, flap porous tips, and landing gear fairings. 

These technologies are considered mature enough to be commensurate with the “N+1” timeframe. 
6 

4.0 Study Engine Designs 

4.1 Engine Trade Space 

The engine design trade space investigated is summarized in Table 2. ADP fan pressure ratios of 1.4, 

1.5, 1.6, and 1.7 were investigated for the direct drive fan engines and 1.3, 1.4, 1.5, and 1.6 for the geared 

fan engines. For the direct drive engines both “high work” and “low work” LPC approaches were 

examined. Since the benefits of the geared fan design are most pronounced when the low-spool is 

required to do a large amount of compression work, the low work, geared fan combination was not 

modeled. As found in the previous work, a variable area fan nozzle was only needed for fan pressure 

ratios of 1.4 and below. Not all combinations of parameters necessarily lead to practical designs. An 

exceptionally large engine diameter is an issue for the geared, FPR=1.3 engine. The large diameter leads 

to a long landing gear given the under-wing engine integration approach assumed for the study. More 

detailed analysis would be necessary to determine if such a long landing gear could be practically 

accommodated. The direct drive, FPR=1.4 design is an extreme case for the direct drive engine 

architecture. Under the design ground rules used for this study, the slow low-spool speed of this engine 

necessitates a large number of LPT stages, leading to an extremely heavy and long engine. These 

“impractical” engines were carried forward through the remainder of the analyses, however, in order to

investigate performance trends as FPR is decreased to the extreme. There may also be practicality issues 

for the FPR=1.6 geared engine. Although the fan gear ratio for the FPR=1.3 engine is 3.6, this ratio 

decreases to only 1.6 for the FPR=1.6 geared engine. Gear ratios close to 1.0 may be difficult to justify 

from a mechanical design standpoint since any weight or performance benefits that might result would 

likely be overwhelmed by the increased design and operational costs associated with the gearbox. 

Table 2. Engine Trade Space 

Engine 

Designation Fan Drive Fan Nozzle ADP FPR OPR LPC PR HPC PR 

Lo-dd-1.4* Direct Variable M0.80/35kft 1.4 42 1.69 17.7 

Lo-dd-1.5 Direct Fixed M0.80/35kft 1.5 42 1.58 17.7

Lo-dd-1.6 Direct Fixed M0.80/35kft 1.6 42 1.48 17.7

Lo-dd-1.7 Direct Fixed M0.80/35kft 1.7 42 1.39 17.7

Hi-dd-1.4* Direct Variable M0.80/35kft 1.4 42 2.50 12.0 

Hi-dd-1.5 Direct Fixed M0.80/35kft 1.5 42 2.33 12.0

Hi-dd-1.6 Direct Fixed M0.80/35kft 1.6 42 2.19 12.0

Hi-dd-1.7 Direct Fixed M0.80/35kft 1.7 42 2.06 12.0

Hi-g-1.3* Geared Variable M0.80/35kft 1.3 42 2.69 12.0 

Hi-g-1.4 Geared Variable M0.80/35kft 1.4 42 2.50 12.0 

Hi-g-1.5 Geared Fixed M0.80/35kft 1.5 42 2.33 12.0 

Hi-g-1.6 Geared Fixed M0.80/35kft 1.6 42 2.19 12.0 

*Design ground rules lead to practicality issues for these cases. 

General characteristics of the low work and high work engine designs are shown in Tables 3 and 4, 

respectively. The counter trends of increasing engine size and weight versus decreasing TSFC as FPR is 

decreased are clearly evident in both the low work and high work designs. For the low work engines, 

TSFC at ADP decreases from 0.525 lb/(lb-h) with a FPR of 1.7 to 0.479 lb/(lb-h) with a FPR of 1.4 (-

9%), reflecting the improvement in propulsive efficiency associated with increasing BPR from 10 to 19.

Note that even at a FPR=1.7, high by today’s standards, the BPR of 10 is much higher than that of current 

engines in this thrust class. The higher BPR for a given FPR compared to current engines is a reflection of 

the significant reduction in core size enabled by the advanced technologies assumed in the engine design. 
7 

The TSFC improvement with lower FPR comes with a significant increase in engine weight. The weight 

of the direct drive, FPR=1.4 engine is 73% more than the direct drive, FPR=1.7 design. The geared fan 

approach included in the high work designs does mitigate the weight penalty associated with low FPR. 

The geared, high work FPR=1.4 engine is only 15% heavier than the direct drive, high work FPR=1.7 

engine, while retaining a 8% TSFC benefit. Even though the geared fan approach decreases the weight 

penalties of low FPR, the issues associated with increased nacelle drag and engine integration remain. 
8 

Table 3. General Characteristics of Low Work Engines 

Lo-dd-1.4* Lo-dd-1.5 Lo-dd-1.6 Lo-dd-1.7 

Fan Drive/Gear Ratio Direct Drive Direct Drive Direct Drive Direct DriveFan Diameter, in 81 74 69 66 Fan Nozzle Geometry Variable Fixed Fixed Fixed Engine+Nacelle Weight, lb 10563 7965 6592 6099 

Nacelle Max Diameter, ft 8.3 7.6 7.1 6.7 

Operating Conditions SLS TOC SLS TOC SLS TOC SLS TOC 

Fan Pressure Ratio 1.3 1.4 1.4 1.5 1.5 1.6 1.6 1.7 

Bypass Ratio 18.4 18.8 15.0 14.8 12.7 12.5 10.9 10.5 

Overall Pressure Ratio 33.1 42 33.8 42 34.4 42 35 42 

Net Thrust, lb 23813 5000 23370 5000 23046 5000 22734 5000 

TSFC, lb/(lb-h) 0.233 0.479 0.253 0.495 0.271 0.510 0.290 0.525 

NOX Emission Index (g/kg) 25.4 8.1 25.8 7.8 26.4 7.6 26.9 7.4 

*Design ground rules lead to practicality issues for this case. 

Table 4. General Characteristics of High Work Engines 

Hi-g-1.3* Hi-g-1.4 Hi-dd-1.4* Hi-g-1.5 Hi-dd-1.5 Hi-g-1.6 Hi-dd-1.6 Hi-dd-1.7 

Fan Drive/Gear Ratio Geared/3.6 Geared/2.6 Direct Drive Geared/2.0 Direct Drive Geared/1.6 Direct Drive Direct Drive Fan Diameter, in 90 80 80 74 74 69 69 66 Fan Nozzle Geometry Variable Variable Variable Fixed Fixed Fixed Fixed Fixed Engine+Nacelle Weight, lb 8736 7401 11292 6626 8821 6252 7276 6451 

Nacelle Max Diameter, ft 9.3 8.2 8.2 7.6 7.6 7.1 7.1 6.7 

Operating Conditions SLS TOC SLS TOC SLS TOC SLS TOC SLS TOC SLS TOC SLS TOC SLS TOC 

Fan Pressure Ratio 1.2 1.3 1.3 1.4 1.3 1.4 1.4 1.5 1.4 1.5 1.5 1.6 1.5 1.6 1.6 1.7 

Bypass Ratio 24.3 24.2 17.6 17.5 17.8 17.7 14.7 14.3 14.8 14.4 12.4 11.8 12.6 12.0 10.8 10.2 

Overall Pressure Ratio 31.7 42 32.9 42 32.9 42 32.7 42 32.7 42 33.1 42 33.1 42 33.6 42 

Net Thrust, lb 26343 5000 24917 5000 24915 5000 23369 5000 23365 5000 22924 5000 22920 5000 22561 5000 

TSFC, lb/(lb-h) 0.204 0.470 0.236 0.486 0.234 0.481 0.257 0.502 0.254 0.497 0.276 0.517 0.273 0.512 0.292 0.527 

NOX Emission Index (g/kg) 23.9 7.8 24.9 7.1 24.9 7.1 24.6 7.3 24.6 7.3 25.0 7.0 24.9 7.0 25.4 6.8 

 *Design ground rules lead to practicality issues for these cases. 
9 

4.2 Engine Comparison 

Variation of engine characteristics with FPR and comparisons to the equivalent “Spiral 2” engines of 

the previous study are shown graphically in Figures 2 through 6. The “Spiral 2” results of the previous 

study are shown as faded lines in the figures and the lowest FPR points are connected with dashed lines to 

indicate the practicality issues with those designs. The relationship between BPR and FPR is shown in 

Figure 2. As FPR is decreased to 1.3, BPR increases to nearly 25, almost 5 times that of the CFM56-7B 

engines currently used on the Boeing 737. Bypass ratios for the new engines are generally slightly higher 

than for the engines of the previous study. The increase in BPR is primarily associated with a reduction in 

total chargeable cooling resulting from the change in LPT cooling philosophy. As the demand for cooling 

air decreases, the core size decreases, increasing the BPR. The removal of the uncooled LPT constraint

greatly reduced the HPT chargeable cooling for the high work designs. In the low work engines, more of

the thermal energy entering the HPT is converted to work to power the HPC (which has a higher pressure

ratio compared to the equivalent high work engine) and less HPT chargeable cooling was required to 

achieve an uncooled LPT. The change in cooling philosophy therefore had a much smaller impact on the 

low work designs. At a FPR of 1.5, the BPR increased 10% for the high work engines and only 3.5% for 

the low work engine. 

Figure 2. Relationship between bypass ratio and design fan pressure ratio, TOC conditions. 

The variation of TSFC with FPR is approximately linear as shown in Figure 3. The efficiency benefits 

of lower FPR are clearly evident with an ~11% decrease in TSFC as FPR is decreased from 1.7 to 1.3. At

a given FPR, the geared fan engine has slightly higher TSFC than the equivalent direct drive engine due 

mainly to gearbox losses, but the difference is only ~1%. The changes in engine modeling assumptions 

for the current study had little impact on the TSFC characteristics of the low work engines, but resulted in 

a downward shift of the TSFC curves for the high work engines. The change in LPT cooling philosophy 

and associated cycle changes as discussed above are primarily responsible for this reduction in TSFC for 

the high work engines. The TSFC penalty of the high work approach seen in the previous study is 

essentially eliminated with the new engines. 

0

5

10

15

20

25

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

TOC Bypass Ratio

Direct Drive, High

Geared, High

Direct Drive, Low

0

5

10

15

20

25

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

TOC Bypass Ratio

Direct Drive, High

Geared, High

Direct Drive, Low
10 

Figure 3. Variation of TSFC with engine type and design fan pressure ratio, TOC conditions. 

The variation of engine size (nacelle diameter) with FPR is shown in Figure 4. The choice of high 

work versus low work, or direct drive versus geared fan has no impact on the engine size. In all cases, as 

FPR is decreased, the fan diameter (and therefore nacelle diameter) must increase to provide the required 

design thrust. There is also essentially no size difference between the engines in this study and the “Spiral 

2” engines in reference 2. 

Figure 4. Variation of nacelle maximum diameter with engine type and design fan pressure ratio. 

Engine weight estimates are presented in Figure 5. For both the direct drive and geared fan engines, 

there is a sizable weight penalty associated with decreasing FPR; however, this penalty is much less 

severe for the geared fan architecture. As FPR is reduced, the fan and overall diameter of the engine 

become larger, leading to higher fan and nacelle weight. This is the primary cause of the increased weight 

of the low FPR geared engines. For the direct drive engines, as the FPR and fan rotational speed decrease, 

0.45

0.47

0.49

0.51

0.53

0.55

0.57

0.59

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

TOC TSFC, lb/(lb-h)

Direct Drive, High

Geared, High

Direct Drive, Low

0.45

0.47

0.49

0.51

0.53

0.55

0.57

0.59

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

TOC TSFC, lb/(lb-h)

Direct Drive, High

Geared, High

Direct Drive, Low

0

2

4

6

8

10

12

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Nacelle Diameter, ft

Direct Drive, High

Geared, High

Direct Drive, Low

0

2

4

6

8

10

12

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Nacelle Diameter, ft

Direct Drive, High

Geared, High

Direct Drive, Low
11 

so does the rotational speed of the directly connected LPT. At a slower speed, more LPT stages are 

needed resulting in an additional weight penalty not present in the geared designs. For the geared engines, 

LPT rotational speed can remain high as FPR is reduced and LPT weight varies little with FPR. The new 

design assumptions had little impact on the weight trends, but did result in an approximately 6% increase 

in engine weight for the direct drive engines. In general this weight increase is due to higher component 

diameters and/or additional stages. For example, an additional LPT stage was required for most of the 

high work, direct drive designs to meet the higher LPT work output of the revised engine cycles while 

keeping LPT loading the same. The geared engines can more readily accommodate these types of changes 

with minimal weight impact through changes in the gear ratio. For example, only a slight (<5%) increase 

in the gear ratio was necessary to accommodate the increase in LPT work, with no changes in the LPT 

loading or number of stages. 

Figure 5. Variation of engine+nacelle weight with engine type and design fan pressure ratio. 

The comparison in Figure 5 is not completely valid because of the difference in the thrust sizing points 

for the current and previously designed engines. Although all the engines represented in Figure 5 have

equal thrust at Mach 0.8, 35,000ft, the sea level static and rolling takeoff thrusts vary. In Figure 6, the 

weight data is presented in terms of engine thrust-to-weight ratio (T/W) at the rolling takeoff condition 

(M=0.25, SL). Although the thrust-to-weight ratios for the current engines are lower at high FPR, the 

T/W curves are flatter than for the engines of reference 2. In other words, with the new design approach 

of using rolling takeoff as a thrust sizing point, the weight penalty to provide equal thrust at rolling 

takeoff as FPR is lowered has been reduced. This results in higher takeoff T/W at low FPR. Since takeoff 

performance was an active constraint in the aircraft sizing analysis of reference 2, this should result in 

better overall aircraft system level performance for the low FPR engines. 

0

2000

4000

6000

8000

10000

12000

14000

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Engine+Nacelle Weight, lb

Direct Drive, High

Geared, High

Direct Drive, Low

0

2000

4000

6000

8000

10000

12000

14000

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Engine+Nacelle Weight, lb

Direct Drive, High

Geared, High

Direct Drive, Low
12 

0

0.5

1

1.5

2

2.5

3

3.5

4

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Engine T/W at M=0.25, SL

Direct Drive, High

Geared, High

Direct Drive, Low

0

0.5

1

1.5

2

2.5

3

3.5

4

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Engine T/W at M=0.25, SL

Direct Drive, High

Geared, High

Direct Drive, Low

Figure 6. Variation of engine thrust-to-weight ratio with engine type and design fan pressure ratio. 

 

5.0 Aircraft Sizing and Performance Results 

It is not clear from the engine characteristics alone what impact the revised engine designs will have

on the conclusions of reference 2. For the low work, direct drive engines, TSFC is approximately the 

same, but weight is higher. For the high work engines there is a decrease in TSFC, but an increase in 

weight, although only a very slight increase for the geared fan engines. Only by combining the engines

with the airframe model can the net impact of these changes on metrics such as block fuel consumption be 

determined. For each of the engine cases, FLOPS (with additional external analyses for wing weight, 

fuselage weight, landing gear size, vertical tail size, and engine-out drag) was used to size/optimize the 

gross weight, wing area, and thrust needed to perform the design mission. Other design parameters such

as wing aspect ratio and wing taper ratio were held fixed at the 737-800 values. Constraints used in wing 

and engine sizing included: wing fuel capacity, approach speed, missed approach climb, second segment 

climb, takeoff field length, and initial cruise altitude capability (expressed as minimum potential rate-ofclimb of 300 ft/min at M=0.8, 35,000ft). Other performance aspects of the configurations (e.g., landing 

field length) were also checked for reasonableness following completion of the sizing/optimization. 

Results of the aircraft sizing analysis are summarized in Table 5 for the low work engines and Table 6

for the high work engines. The numbers in bold indicate the “best” value for that particular parameter. 

Results for a 1998 entry-into-service (EIS) technology baseline vehicle (similar to a 737-800 with 

CFM56-7B engines, but sized to perform the study design mission) are shown in the tables for 

comparison. For the low work, direct drive engines, block fuel consumption is minimized with a design 

FPR of 1.6. The BPR for this engine is 12.5 at ADP conditions. Note that the cruise range factor, which is 

an approximate measure of combined aerodynamic and propulsive efficiency, increases at lower FPR. 

However, this higher efficiency is insufficient to offset the increase in weight that occurs for the lower 

FPR engines. This clearly illustrates the system level trade-off that occurs for lower FPR engines, the 

benefit of higher efficiency versus the penalty of higher weight. Although the minimum fuel consumption 

case is at FPR=1.6, the minimum gross weight occurs at FPR=1.7. In other words, the decrease in engine

weight for that configuration is sufficient to offset the increase in fuel weight to arrive at a lower total 
13 

gross weight. Lowest total NOX emissions (referred to as “block NOX” in the table) also occur with a 

FPR=1.7 design. Landing-takeoff cycle (LTO) NOX emissions are lowest for the FPR=1.4 engine; but, the 

variation in LTO NOX is relatively small among all the cases. Clearly, identifying a “best” engine design 

depends on the metric of interest. Ultimately the primary metric is life cycle cost, and historically gross 

weight has been used as a surrogate for life cycle cost in aircraft design and optimization. However, 

recent increases in fuel cost have made fuel consumption a more important factor in life cycle cost. It may 

no longer be valid to assume the lowest gross weight configuration has the lowest life cycle cost. 

Sizing results for the high work engine cases are given in Table 6. Both geared and direct drive fan 

approaches were considered for the high work engines. The lowest block fuel consumption occurs for the

geared, FPR=1.5 engine case. The BPR of this engine is 14.3 at ADP conditions. As with the low work 

engines, there is a trade-off between the efficiency associated with lower FPR and the increase in engine 

weight. The FPR=1.3 and FPR=1.4 geared fan cases have higher cruise efficiency, but also higher block 

fuel consumption. The geared, FPR=1.6 case provides the lowest total NOX emissions and lowest ramp 

weight. The lowest LTO NOX emissions occur at the opposite end of the fan pressure ratio spectrum, at 

FPR=1.3. The geared fan system is able to mitigate to some extent the penalties associated with 

decreasing FPR and increasing BPR. This benefit can be seen by comparing the FPR=1.4 results in Table 

6 for the two different fan drive approaches. The aircraft with the geared fan engine has a lower empty 

weight, lower ramp weight, lower block fuel consumption, lower total NOX and lower LTO NOX

emissions. But, even in the case of geared designs, minimum block fuel occurs at a FPR of 1.5, not at 

lower FPR where cruise efficiency is higher. 

Most of the aircraft in Tables 5 and 6, have lower wing loading than is typical for a 737-800 like 

design. For the chosen design mission and constraints, takeoff field length is the primary sizing 

constraint, which can be met by a range of different engine and wing sizes. The penalty of increasing 

wing size is diminished somewhat for the ASAT configurations relative to current designs due to the use 

of composite materials. As a result, for the ASAT designs the preferred (lower gross weight) approach to 

meeting the takeoff requirement tends to be a larger wing (low wing loading) rather than a larger engine 

(high thrust-to-weight). This lower wing loading also reduces approach speed, benefiting approach noise 

levels. 
14 

Table 5. Aircraft Sizing Results for Low Work Engines (162 Passenger, 3250 nm Design Mission) 

1998 Tech. 

Baseline 

Lo-dd-1.4* 

(BPR ~ 19) 

Lo-dd-1.5 

(BPR ~ 15) 

Lo-dd-1.6 

(BPR ~ 13) 

Lo-dd-1.7 

(BPR ~ 11) 

OEW, lb 94700 97450 87300 82800 81400 Mission Fuel, lb 50350 38750 36700 36400 37000 Payload Weight, lb 32400 32400 32400 32400 32400 Ramp Weight, lb 177550 168600 156400 151600 150800 Wing Area, ft2 1470 1450 1390 1330 1320 W/S, lb/ft2 121 116 112 114 114 Thrust(SLS), lb 26100 26100 23600 22900 22650 T/W (takeoff) 0.294 0.310 0.302 0.303 0.300 Takeoff field length, ft 7000 7000 7000 7000 7000 Landing field length, ft 6020 5850 5700 5740 5750 ~Cruise Range Factor  V*(L/D)/TSFC, nm 12450 15600 15300 15000 14700 Block Fuel, lb 42600 32900 31100 30800 31250 Block NOX, lb 555 245 226 218 216 LTO NOX, lb per cycle 22.2 10.7 11.0 11.1 11.8 

*Design ground rules lead to practicality issues for this case. 
15 

Table 6. Aircraft Sizing Results for High Work Engines (162 Passenger, 3250 nm Design Mission) 

1998 Tech. 

Baseline 

Hi-g-1.3* 

(BPR ~ 24) 

Hi-g-1.4 

(BPR ~ 18) 

Hi-dd-1.4* 

(BPR ~ 18) 

Hi-g-1.5 

(BPR ~ 15) 

Hi-dd-1.5 

(BPR ~ 15) 

Hi-g-1.6 

(BPR ~ 12) 

Hi-dd-1.6 

(BPR ~ 12) 

Hi-dd-1.7 

(BPR ~ 11) 

OEW, lb 94700 90200 85300 101050 83000 90650 81950 85150 82750Mission Fuel, lb 50350 36400 36100 40800 35850 38000 36600 37300 37550 Payload Weight, lb 32400 32400 32400 32400 32400 32400 32400 32400 32400 Ramp Weight, lb 177550 158950 153800 174200 151200 161000 151000 154800 152700 Wing Area, ft2 1470 1340 1340 1460 1340 1420 1325 1380 1350 W/S, lb/ft2 121 119 114 120 113 114 114 112 113 Thrust(SLS), lb 26100 27650 24800 28350 23200 24400 22900 23100 22600 T/W (takeoff) 0.294 0.348 0.323 0.325 0.306 0.303 0.303 0.298 0.296 Takeoff field length, ft 7000 7000 7000 7000 7000 7000 7000 7000 7000 Landing field length, ft 6020 5940 5780 5970 5730 5750 5760 5700 5720 ~Cruise Range Factor  V*(L/D)/TSFC, nm 12450 15600 15300 15300 15100 15200 14800 14900 14600 Block Fuel, lb 42600 30900 30600 34550 30400 32200 31000 31600 31800 Block NOX, lb 555 226 215 242 210 224 209 215 210LTO NOX, lb per cycle 22.2 9.3 10.0 11.3 10.0 10.4 10.6 10.6 11.2 

 *Design ground rules lead to practicality issues for these cases. 
16 

The weight, fuel consumption, and NOX results are presented graphically in Figures 7 through 13, 

overlaid on the reference 2 results. Aircraft empty weight results are compared in Figure 7. Consistent 

with the trends in engine weight, aircraft empty weight increases as FPR is decreased. Also, the increase 

in engine weight relative to reference 2 is reflected in the generally higher aircraft empty weights for the 

current analysis. Trends are similar to the previous results, although the curves are flatter and empty 

weight is now less sensitive to FPR than before. This is particularly true for the high work, geared engine 

cases in which the previous empty weight penalty of ~20,000 lb associated with decreasing FPR from 1.6

to 1.3 has been reduced to only 8,250 lb. 

Figure 7. Variation of aircraft empty weight with engine type and design fan pressure ratio. 

The impacts of the engine design changes on ramp weight are similar to the impacts on empty weight. 

The results presented in Figure 8 indicate higher ramp weights for FPRs of 1.6 and 1.7, similar ramp 

weights for FPR=1.5, and lower ramp weights for the low FPRs of 1.3 and 1.4. Although the sensitivity of 

ramp weight to FPR has been reduced, the lowest ramp weight cases still occur for the highest fan 

pressure ratios analyzed. 

60000

70000

80000

90000

100000

110000

120000

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Operating Empty Weight

Direct Drive, High

Geared, High

Direct Drive, Low

60000

70000

80000

90000

100000

110000

120000

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Operating Empty Weight

Direct Drive, High

Geared, High

Direct Drive, Low
17 

Figure 8. Variation of aircraft ramp weight with engine type and design fan pressure ratio. 

Fuel consumption results are presented in Figure 9. For the direct drive engines, the same trends are 

observed as in the results of reference 2. For both low work and high work cases, minimum block fuel 

consumption occurs with a FPR of 1.6. The new engines result in slightly higher fuel consumption for 

FPR of 1.5 to 1.7 and lower fuel consumption for FPR of 1.4. The engine design changes had greater 

impacts on the high work geared engine cases. The high work geared cases benefited from both a 

decrease in TSFC (Figure 3) and an increase in thrust-to-weight at the rolling takeoff condition (Figure 6). 

Fuel consumption is reduced in all cases, and reducing FPR has a much smaller negative affect on the fuel 

consumption due to lower engine weight penalties for low FPR (flatter curve in Figure 6). Note, however, 

that although the variation of fuel consumption with FPR is fairly small for the geared, high work 

engines, the minimum block fuel still occurs at a FPR of 1.5, as was also found in reference 2. The engine 

design changes have not changed the optimum FPR for fuel consumption for a given engine architecture, 

but the overall minimum fuel case has changed. Previously the overall minimum fuel consumption 

occurred for the low work, direct drive fan, FPR=1.6 case. With the changes in engine design, the 

minimum fuel consumption now occurs for the high work, geared fan, FPR=1.5 case. The “cross over” 

FPR below which the geared fan approach results in lower fuel consumption than the low work, direct 

drive approach has also moved from FPR=1.5 to FRP~1.55. Although the geared fan, FPR=1.5 case 

results in the lowest fuel consumption of all the cases, the low work, direct drive FPR=1.6 case has a fuel 

consumption that is only 1.3% higher. Given the high level nature of this analysis, these two cases can be 

considered essentially equivalent from a fuel consumption perspective. The lower fan pressure ratio of the 

geared fan, FPR=1.5 case leads to lower noise, however, as will be discussed later. 

130000

140000

150000

160000

170000

180000

190000

200000

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Ramp Weight

Direct Drive, High

Geared, High

Direct Drive, Low

130000

140000

150000

160000

170000

180000

190000

200000

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Ramp Weight

Direct Drive, High

Geared, High

Direct Drive, Low
18 

Figure 9. Variation of block fuel with engine type and design fan pressure ratio. 

For total NOX, shown in Figure 10, the minimum NOX emissions occur at the highest FPR analyzed 

for all the engine architectures. Results are similar between the high work and low work direct drive

cases, while the high work geared fan engines result in lower total NOX emissions, particularly at low 

FPRs. Compared to the results of reference 2, the NOX is higher for high FPR cases and lower for low 

FPR cases, reflecting the same “flattening” of the curves as seen in the other metrics. 

The landing-takeoff cycle NOX (LTO NOX) can be expressed in a number of different ways. LTO NOX

is regulated as an engine parameter, “Dp/Foo” where Dp is the grams of NOX emitted over a standard LTO 

cycle (by a single, uninstalled engine) and Foo is the rated output at SLS conditions in kilonewtons. This 

parameter is defined by the International Civil Aviation Organization (ICAO) and used in FAR Part 34 

for engine certification (ref. 22). The results for Dp/Foo are presented in Figure 11; results are similar to

those obtained in reference 2. There is a consistent downward trend in Dp/Foo with decreasing fan 

pressure ratio. Another way to view LTO NOX is the margin relative to the current CAEP6 regulatory 

limit, as shown in Figure 12. CAEP6 refers to the limits adopted at the sixth meeting of the ICAO Council 

Committee on Aviation Environmental Protection (CAEP), held in 2004. Not only does lower FPR 

decrease Dp/Foo, it also increases the margin relative to the regulation. Note that all of the engines are 

predicted to meet or exceed NASA’s CAEP6-60% goal for the “N+1” timeframe. The Dp/Foo emission 

parameter alone does not account for differences in engine weight and performance which can lead to 

differences in the required thrust level (Foo) when the engine is integrated into an overall aircraft design. 

Lower Dp/Foo does not necessarily result in lower total LTO NOX emissions. The estimated NOX per LTO 

cycle is compared in Figure 13. NOX per LTO cycle has been estimated by multiplying the ICAO Dp/Foo

parameter by the total engine thrust. The trends shown in Figure 13, are not as consistent as the other 

metrics. High FPR certainly leads to higher NOX, but between FPR=1.3 and 1.5 the variation with FPR is 

not consistent. The NOX per LTO cycle results depend on a combination of the engine characteristics 

(Dp/Foo) and the aircraft sizing results (e.g., trade between engine thrust and wing area necessary to meet 

takeoff performance); therefore, they exhibit more variability. 

30000

31000

32000

33000

34000

35000

36000

37000

38000

39000

40000

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Block Fuel

Direct Drive, High

Geared, High

Direct Drive, Low

30000

31000

32000

33000

34000

35000

36000

37000

38000

39000

40000

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Block Fuel

Direct Drive, High

Geared, High

Direct Drive, Low
19 

0

5

10

15

20

25

30

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

LTO NOX, g/kN

Direct Drive, High

Geared, High

Direct Drive, Low

Figure 10. Variation of total mission NOX emissions with engine type and design fan pressure ratio. 

0

5

10

15

20

25

30

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

LTO NOX, g/kN

Direct Drive, High

Geared, High

Direct Drive, Low

Figure 11. Variation of LTO NOX emissions with engine type and design fan pressure ratio (Dp/Foo). 

100

120

140

160

180

200

220

240

260

280

300

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Block NOX, lb

Direct Drive, High

Geared, High

Direct Drive, Low

100

120

140

160

180

200

220

240

260

280

300

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Block NOX, lb

Direct Drive, High

Geared, High

Direct Drive, Low
20 

-80%

-70%

-60%

-50%

-40%

-30%

-20%

-10%

0%

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

LTO NOX, relative to CAEP6

Direct Drive, High

Geared, High

Direct Drive, Low

-80%

-70%

-60%

-50%

-40%

-30%

-20%

-10%

0%

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

LTO NOX, relative to CAEP6

Direct Drive, High

Geared, High

Direct Drive, Low

Figure 12. LTO NOX margin relative to CAEP6 regulatory limit. 

Figure 13. Variation of NOX emissions per LTO cycle with engine type and design fan pressure ratio. 

Overall fuel, weight, and NOX benefits of the ASAT engine and airframe technologies are shown in 

Figures 14 through 17. Regardless of the engine design chosen, fuel burn reduction relative to the 1998 

technology baseline is in the range of 25-30% (Figure 14), with the largest reduction of 28.6% for the

high work, geared, FPR=1.5 engine design. These significant reductions in block fuel are a result of the 

combination of advanced airframe and engine technology assumptions, in addition to the engine cycle 

changes. Reductions in ramp weight on the order of 15% are possible as well, as shown in Figure 15. The 

advanced combustor technology assumed for the engines in this study results in large reductions in both 

0

2

4

6

8

10

12

14

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

LTO NOX, lb

Direct Drive, High

Geared, High

Direct Drive, Low

0

2

4

6

8

10

12

14

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

LTO NOX, lb

Direct Drive, High

Geared, High

Direct Drive, Low
21 

total NOX and LTO NOX. Reductions of 60% in total NOX and 50-60% in LTO NOX are estimated for the 

advanced ASAT vehicles as shown in Figure 16 and Figure 17, respectively. 

Figure 14. Potential block fuel reductions from application of ASAT technologies. 

Figure 15. Potential ramp weight reductions from application of ASAT technologies. 

-16%

-14%

-12%

-10%

-8%

-6%

-4%

-2%

0%

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Change in Ramp Weight*

Direct Drive, High

Geared, High

Direct Drive, Low

* Baseline is 1998 EIS Technology Airframe and Engine 

-30%

-25%

-20%

-15%

-10%

-5%

0%

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Change in Block Fuel*

Direct Drive, High

Geared, High

Direct Drive, Low

* Baseline is 1998 EIS Technology Airframe and Engine 
22 

-80%

-70%

-60%

-50%

-40%

-30%

-20%

-10%

0%

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Change in Block NOX*

Direct Drive, High

Geared, High

Direct Drive, Low

* Baseline is 1998 EIS Technology Airframe and Engine 

Figure 16. Potential total NOX reductions from application of ASAT technologies. 

Figure 17. Potential LTO NOX reductions from application of ASAT technologies. 

Noise benefits for the ASAT configurations are shown in Figure 18. Often certification noise is 

expressed in terms of the algebraic sum of the lateral, flyover, and approach noise levels (“cumulative 

noise”). The cumulative noise reductions for the study configurations relative to the 737-800 predicted 

value are presented in Figure 18. The noise reduction estimates from reference 2 are also shown for 

comparison. As found previously in reference 2, the noise results are dominated by FPR. Fan drive 

approach and compression work split have little impact on the noise characteristics. The positive impact 

of low FPR on noise is clearly evident. The cumulative noise reduction achieved by reducing FPR from 

1.7 to 1.3 is over 21 EPNdB; despite the higher aircraft weight and engine thrust associated with FPR=1.3 

-60%

-50%

-40%

-30%

-20%

-10%

0%

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Change in NOX per LTO*

Direct Drive, High

Geared, High

Direct Drive, Low

* Baseline is 1998 EIS Technology Airframe and Engine 
23 

-35.0

-30.0

-25.0

-20.0

-15.0

-10.0

-5.0

0.0

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Change in Cumulative EPNL, EPNdB

Direct Drive, High

Geared, High

Direct Drive, Low

* Baseline is predicted 737-800 noise level

case. The maximum reduction relative to the 737-800 baseline realized for a practical design is 25.3 

EPNdB cumulative (geared, FPR=1.4 case). This reduction is the result of both engine cycle changes and

noise reduction technologies. At FPR=1.3, the cumulative noise reduction is over 30 EPNdB. The 

FPR=1.3 case in this study results in an aircraft design with landing gear length issues. However, there 

may be other integration approaches which would enable the additional noise benefit of the FPR=1.3 

engine to be realized in a practical aircraft design. It was previously noted that the fuel consumption 

characteristics of the minimum fuel case (low work, geared fan, FPR=1.5) could be nearly achieved with

the low work, direct drive fan, FPR=1.6 case. The geared, FPR=1.5 approach, however, results in 6.7 

EPNdB lower cumulative noise than the direct drive, FRP=1.6 case, as can be observed in Figure 18. 

Although the geared engines do not offer a significant fuel consumption benefit over the direct drive 

engines, they do offer the potential for significantly lower noise with equivalent fuel consumption. 

Compared to the results in reference 2, the changes in engine design have generally resulted in a slight 

decrease in cumulative noise of approximately 2 EPNdB or less. 

-35.0

-30.0

-25.0

-20.0

-15.0

-10.0

-5.0

0.0

1.2 1.3 1.4 1.5 1.6 1.7 1.8

Top-of-Climb Fan Pressure Ratio

Change in Cumulative EPNL, EPNdB

Direct Drive, High

Geared, High

Direct Drive, Low

* Baseline is predicted 737-800 noise level

Figure 18. Potential noise reductions from application of ASAT technologies. 

6.0 Design Trade-Offs 

Since no single engine provides the best performance across all of the metrics, there are a series of 

design trade-offs that must be accepted. For example, choosing a very low FPR because of the noise 

benefits incurs an increase in aircraft weight compared to what could be achieved with a higher FPR 

design. The magnitudes of these various trade-offs are presented in Table 7. Similar analysis was also

performed in reference 2 and those results are shown in parentheses in Table 7. For each metric column, 

the minimum value is used as the baseline for the remaining cases. The values in the remaining rows 

indicate the magnitude of the penalty incurred from selecting that particular engine, relative to the best 

possible result. For example, the direct drive, low work, FPR=1.5 case has 3.7% higher ramp weight than 

the minimum (direct drive, low work LPC, FPR=1.7); 2.3% greater block fuel consumption than the 

minimum (geared, high work LPC, FPR=1.5); 8.4% greater block NOX than the minimum (geared, high 

work LPC, FPR=1.6); 10.6% greater LTO NOX than the minimum (geared, high work LPC, FPR=1.4);

and 4.3 EPNdB higher cumulative noise than the minimum (geared, high work LPC, FPR=1.4). (The 

geared, FPR=1.3 and direct drive, FPR=1.4 cases were not included in the trade-off analysis because of
24 

the practicality issues with those designs.) 

Table 7. Weight, Fuel, Emissions, and Noise Trade-Offs 

Ramp Weight Block Fuel Block NOX LTO NOX Cum. EPNdB

High, Geared, FPR=1.4 +2.0 %

(+7.3%)

+0.5%

(+5.9%)

+2.7%

(+15.8%)

Minimum

(+2.3%)

Minimum

(Minimum)

High, Geared, FPR=1.5

+0.3%

(+2.3%)

Minimum

(+1.6%)

+0.5%

(+5.8%)

+0.5%

(Minimum)

+3.7

(+5.3)

Low, Direct, FPR=1.5 +3.7%

(+4.3%)

+2.3%

(+1.1%)

+8.4%

(+12.8%)

+10.6%

(+0.8%)

+4.3

(+5.4)

High, Direct, FPR=1.5 +6.8%

(+7.8%)

+6.0%

(+6.2%)

+7.3%

(+11.2%)

+4.8%

(+2.8%)

+4.4

(+5.8)

High, Geared, FPR=1.6

+0.1%

(+1.3%)

+2.0%

(+2.7%)

Minimum

(+3.1%)

+6.9%

(+6.1%)

+10.3

(+12.4)

Low, Direct, FPR=1.6 +0.5%

(+1.0%)

+1.2%

(Minimum)

+4.5%

(+8.1%)

+11.5%

(+5.1%)

+10.4

(+11.9)

High, Direct, FPR=1.6 +2.6%

(+3.1%)

+3.9%

(+3.7%)

+3.0%

(+4.7%)

+6.9%

(+5.9%)

+10.5

(+12.5)

Low, Direct, FPR=1.7

Minimum

(Minimum)

+2.8%

(+1.3%)

+3.4%

(+1.2%)

+18.9%

(+14.3%)

+16.1

(+16.7)

High, Direct, FPR=1.7 +1.2%

(+1.6%)

+4.5%

(+4.4%)

+0.5%

(Minimum)

+12.7%

(+13.4%)

+15.8

(+17.1)

Ramp Weight Block Fuel Block NOX LTO NOX Cum. EPNdB

High, Geared, FPR=1.4 +2.0 %

(+7.3%)

+0.5%

(+5.9%)

+2.7%

(+15.8%)

Minimum

(+2.3%)

Minimum

(Minimum)

High, Geared, FPR=1.5

+0.3%

(+2.3%)

Minimum

(+1.6%)

+0.5%

(+5.8%)

+0.5%

(Minimum)

+3.7

(+5.3)

Low, Direct, FPR=1.5 +3.7%

(+4.3%)

+2.3%

(+1.1%)

+8.4%

(+12.8%)

+10.6%

(+0.8%)

+4.3

(+5.4)

High, Direct, FPR=1.5 +6.8%

(+7.8%)

+6.0%

(+6.2%)

+7.3%

(+11.2%)

+4.8%

(+2.8%)

+4.4

(+5.8)

High, Geared, FPR=1.6

+0.1%

(+1.3%)

+2.0%

(+2.7%)

Minimum

(+3.1%)

+6.9%

(+6.1%)

+10.3

(+12.4)

Low, Direct, FPR=1.6 +0.5%

(+1.0%)

+1.2%

(Minimum)

+4.5%

(+8.1%)

+11.5%

(+5.1%)

+10.4

(+11.9)

High, Direct, FPR=1.6 +2.6%

(+3.1%)

+3.9%

(+3.7%)

+3.0%

(+4.7%)

+6.9%

(+5.9%)

+10.5

(+12.5)

Low, Direct, FPR=1.7

Minimum

(Minimum)

+2.8%

(+1.3%)

+3.4%

(+1.2%)

+18.9%

(+14.3%)

+16.1

(+16.7)

High, Direct, FPR=1.7 +1.2%

(+1.6%)

+4.5%

(+4.4%)

+0.5%

(Minimum)

+12.7%

(+13.4%)

+15.8

(+17.1)

For ramp weight, the minimum case has not changed compared to the results from reference 2, but the 

ramp weight penalties of the other cases have been reduced. For example, previously the ramp weight 

penalty of the high work, geared fan, FPR=1.4 case was 7.3%, versus only 2.0% with the revised engine 

designs. As noted previously, the minimum block fuel case did change from the low work, direct drive, 

FPR=1.6 case to the high work, geared, FPR=1.5 case. The fuel consumption penalty of the FRP=1.4, 

geared engine case was greatly reduced from 5.9% to only 0.5%. The block NOX penalty of the low FPR, 

geared designs was also greatly reduced, from 15.8% to only 2.7% at FPR=1.4. The minimum noise case 

continues to be the geared, FPR=1.4 case, but there is a slight reduction in the noise penalties of the other 

cases. 

With improved performance of the low FPR engines resulting from the engine design changes 

implemented in this study, there is less of a trade-off between aircraft noise and fuel consumption than 

found previously in reference 2. In reference 2, pursuing minimum fuel consumption resulted in a 12 

EPNdB penalty in cumulative noise, whereas pursuing minimum noise resulted in a 6% penalty in fuel 

consumption. For the current study, pursuing minimum fuel consumption (which now occurs at a lower 

FPR) only results in a 4 EPNdB penalty in noise and pursuing minimum noise only results in a 0.5% 

penalty in fuel consumption. 

Which design is better overall depends, in part, on the relative value of low noise, low fuel 

consumption, and low NOX emissions to the airlines, which in turn depends on external factors such as 

fuel cost, airport noise restrictions, and government regulations. Selecting a balanced engine design that 

performs well in all categories is somewhat subjective; however, the geared fan, high work, FPR=1.5 case 

seems to offer the best overall performance. It has relatively good weight, fuel consumption, noise, and 

emissions results. Bypass ratio for this engine is in the 14 range, at the lower end of what is usually 

considered “ultra-high bypass ratio.” If noise were the dominant consideration, a lower FPR would offer 

reduced noise with small weight, fuel, and NOX penalties. The emergence of a high work, geared, 
25 

FPR=1.5 engine as the best balanced design is a function of the approach and assumptions in this study. 

Changes in engine or airframe design rules and technology assumptions could lead to a different result. 

7.0 Concluding Remarks 

A comprehensive exploration of the turbofan engine design space for an advanced technology singleaisle transport was previously documented in reference 2. Through the course of that study and in a 

subsequent evaluation of the approach and results, a number of enhancements to the engine design ground 

rules and assumptions were identified. A follow-on effort was initiated to investigate the impacts of these 

changes on the study results. 

The fundamental conclusions of the prior study have not changed with the changes to the engine 

designs. With the ground rules, architectures, and assumptions used in this study: empty weight and ramp 

weight (often surrogate indicators of cost) are minimized with high FPR; block fuel consumption is 

minimized with a FPR of 1.5-1.6; block NOX emissions are minimized with high FPR; and LTO NOX and 

certification noise are minimized with FPR as low as possible. The best compromise FPR appears to be 

~1.5, having good performance across all the metrics of interest (ramp weight, fuel consumption, 

emissions, and noise). Relative to 1998 EIS technology, the advanced configurations have the potential

for significant benefits: up to 29% reduction in fuel consumption, 50-60% reductions in NOX emissions, 

and greater than 25 EPNdB cumulative noise reduction. 

The engine design changes did change the relative performance of direct drive and geared fan engines 

and the fuel consumption and weight penalties of low FPR. Because of a relative improvement in the 

geared fan cases, the point at which gearing becomes beneficial moved to higher FPR. In the previous 

study, the high work, geared fan and low work, direct drive fan FPR=1.5 engines resulted in similar 

aircraft characteristics. In the current study, the FPR=1.5 geared fan engine is noticeably better than the 

low work, direct drive engine at that FPR. The primary reasons for this shift are the higher relative 

weights of the low work, direct drive engines and the elimination of unnecessary penalties on the high

work engines caused by the previous LPT cooling philosophy. Even though the engine design changes 

did not change the optimum FPR for each of the metrics for a given engine architecture, it did change the 

penalties associated with being away from that optimum. In particular, the penalties of low FPR were 

reduced significantly for the high work, geared fan architecture. This enables lower FPR (and 

correspondingly lower aircraft noise) to be implemented with a minimal fuel consumption penalty. 
26 

8.0 References 

1. Wall, R,; Kingsley-Jones, M.; Norris, G.; Mecham, M.; Warwick, G.; and Shannon, D.: Leap of Faith – Cseries 

killer or not, Airbus launches A320NEO but keeps an eye on what is next. Aviation Week & Space Technology, 

December 6, 2010. pp. 24-28. 

2. Guynn, M. D.; Berton, J. J.; Fisher, K. L.; Haller, W. J.; Tong, M. T.; and Thurman, D. R.: Engine Concept 

Study for an Advanced Single-Aisle Transport. NASA TM-2009-215784, August 2009. 

3. Lytle, J.K.: The Numerical Propulsion System Simulation: An Overview. NASA TM-2000-209915, June 2000. 

4. NPSS User Guide Software Release: NPSS_1.6.3 AL. 

5. NPSS Reference Sheets Software Release: NPSS_1.6.3 AL. 

6. Onat, E.; and Klees, G.: A Method to Estimate Weight and Dimensions of Large and Small Gas Turbine 

Engines. NASA CR 159481, 1979. 

7. 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. 

8. 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. 

9. Engine Comparison. http://www.aviationweek.com/media/images/awst_images/large/Leap-X_PW1000G.html.

Accessed 7/30/2010. 

10. McCullers, L.: Aircraft Configuration Optimization Including Optimized Flight Profiles. Proceedings of the 

Symposium on Recent Experiences in Multidisciplinary Analysis and Optimization, NASA CP 2327, April 

1984. 

11. Ardema, M. D.; Chambers, M. C.; Patron, A. P; Hahn, A. S.; Miura, H.; and Moore, M. D.: Analytical Fuselage 

and Wing Weight Estimation of Transport Aircraft. NASA TM 110392, May 1996. 

12. 787 Dreamliner Program Fact Sheet. http://www.boeing.com/commercial/787family/programfacts.html. 

Accessed 4/9/2007. 

13. 737-800 Technical Characteristics. http://www.boeing.com/commercial/737family/pf/pf_800tech.html. 

Accessed 4/9/2007. 

14. Gillian, Ronnie E.: Aircraft Noise Prediction Program User’s Manual. NASA TM-84486, 1983. 

15. Zorumski, William E.: Aircraft Noise Prediction Program Theoretical Manual. NASA TM-83199, 1981, Parts 

1 and 2 (Currently maintained at NASA LaRC by the ANOPP team in electronic format and provided upon 

request; Latest revision: December 2005). 

16. Code of Federal Regulations, Title 14, Chapter I, Part 36. Noise standards: Aircraft type and airworthiness 

certification.
27 

17. Stone, J.R.; Krejsa, E.A.; Clark, B.J.; and Berton, J.J.: Jet Noise Modeling for Suppressed and Unsuppressed 

Aircraft in Simulated Flight. NASA TM-2009-215524, 2009. 

18. Janardan, B.A.; Hoff, G.E.; Barter, J.W.; Martens, S.; Gliebe, P.R.; Mengle, V.; and Dalton, W.N.: AST Critical 

Propulsion and Noise Reduction Technologies for Future Commercial Subsonic Engines – Separate-Flow 

Exhaust System Noise Reduction Evaluation. Final Report: NAS3–27720, Area of Interest 14.3, General 

Electric Report R98AEB152, May 1998. 

19. Neubert, R.; Bock, L.; Malmborg, E.; and Owen-Peer, W.: Advanced Low Noise Research Fan Stage Design. 

NASA CR 97-206308, 1997. 

20. Sutliff, D.L.; Jones, M.G.; Hartley, T.C.: Attenuation of FJ44 Turbofan Engine Noise with a Foam-Metal Liner 

Installed Over-the-Rotor. 15th AIAA/CEAS Aeroacoustics Conference, AIAA-2009-3141, 2009. 

21. Jones, M.; Parrott, T.; Sutliff, D.; Hughes, C.: Assessment of Soft Vane and Metal Foam Engine Noise

Reduction Concepts. 15th AIAA/CEAS Aeroacoustics Conference, AIAA-2009-3142, 2009. 

22. Code of Federal Regulations, Title 14, Chapter I, Part 34. Fuel Venting and Exhaust Emission Requirements for 

Turbine Engine Powered Airplanes.
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Refined Exploration of Turbofan Design Options for an Advanced 

Single-Aisle Transport 

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Guynn, Mark D.; Berton, Jeffrey J.; Fisher, Kenneth L.; Haller, William J.; 

Tong, Michael T.; Thurman, Douglas R,

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

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14. ABSTRACT

A comprehensive exploration of the turbofan engine design space for an advanced technology single-aisle transport (737/A320 class aircraft) 

was conducted previously by the authors and is documented in a prior report. Through the course of that study and in a subsequent 

evaluation of the approach and results, a number of enhancements to the engine design ground rules and assumptions were identified. A 

follow-on effort was initiated to investigate the impacts of these changes on the original study results. The fundamental conclusions of the 

prior study were found to still be valid with the revised engine designs. The most significant impact of the design changes was a reduction in 

the aircraft weight and block fuel penalties incurred with low fan pressure ratio, ultra-high bypass ratio designs. This enables lower noise 

levels to be pursued (through lower fan pressure ratio) with minor negative impacts on aircraft weight and fuel efficiency. Regardless of the 

engine design selected, the results of this study indicate the potential for the advanced aircraft to realize substantial improvements in fuel 

efficiency, emissions, and noise compared to the current vehicles in this size class. 

15. SUBJECT TERMS

Aircraft Design; Aircraft Noise; Engine Airframe Integration; Engine Design; Propulsion System Performance; Turbofan 

Engines

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