August 2010 

NASA/CR-2010-216842 

N+2 Supersonic Concept Development and 

Systems Integration 

Harry R. Welge, John Bonet, Todd Magee, Daniel Chen, Steve Hollowell, Aaron Kutzmann, 

Alan Mortlock, and Josh Stengle 

Boeing Research & Technology, Huntington Beach, California 

Chet Nelson, Eric Adamson, Steve Baughcum, Robert T. Britt, and Gregory Miller 

Boeing Commercial Airplanes, Seattle, Washington 

Jimmy Tai 

Georgia Institute of Technology, Atlanta, Georgia 
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Hampton, Virginia 23681-2199 under Contract NNL08AA16B 

 

August 2010 

NASA/CR-2010-216842 

N+2 Supersonic Concept Development and 

Systems Integration 

Harry R. Welge, John Bonet, Todd Magee, Daniel Chen, Steve Hollowell, Aaron Kutzmann, 

Alan Mortlock, and Josh Stengle 

Boeing Research & Technology, Huntington Beach, California 

Chet Nelson, Eric Adamson, Steve Baughcum, Robert T. Britt, and Gregory Miller 

Boeing Commercial Airplanes, Seattle, Washington 

Jimmy Tai 

Georgia Institute of Technology, Atlanta, Georgia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 
5 

Table of Contents 

0.0 ABSTRACT ............................................................................................................................................ 8

0.1 EXECUTIVE SUMMARY .................................................................................................................... 9

1.0 MARKET STUDY AND GOALS ....................................................................................................... 11

1.1 TIMING ................................................................................................................................................. 11

1.2 GENERAL MARKET & REGULATION FORECAST ...................................................................................... 11

1.3 INITIAL MARKET/DESIGN REQUIREMENTS & OBJECTIVES ...................................................................... 14

2.0 CONCEPT DEVELOPMENT ............................................................................................................ 22

2.1 INITIAL CONCEPTS ................................................................................................................................ 22

2.2 100 PASSENGER HIGH PERFORMANCE TRANSPORT ................................................................................ 24

2.3 THIRTY (30) PASSENGER, LOWER BOOM TRANSPORT ............................................................................. 32

3.0 PROPULSION DEVELOPMENT .....................................................................................................108

3.1 TRADE SPACE STUDY ENGINES .............................................................................................................108

3.2 RR NPSS DEVELOPMENT AND CYCLE TRADE PARAMETERS ..................................................................108

3.3 THE COUPLED NPSS/MDA TOOL ........................................................................................................113

3.4 072B OPTIMIZATION ............................................................................................................................114

3.5 076E OPTIMIZATION ............................................................................................................................116

3.6 SELECTED CYCLE PARAMETERS ...........................................................................................................117

4.0 NON-PROPRIETARY SMALL SUPERSONIC AIRCRAFT ........................................................148

4.1 CONCEPT SELECTION ..........................................................................................................................148

4.2 DESCRIPTION OF THE 076E .................................................................................................................148

4.3 SONIC BOOM DISCUSSION ....................................................................................................................150

4.4 GOAL COMPLIANCE .............................................................................................................................150

4.5 PROPULSION .......................................................................................................................................151

5.0 SECOND ITERATION ANALYSIS OF BASELINE CONCEPT ..................................................171

5.1 STRUCTURAL ANALYSIS (OPTION 1) ......................................................................................................171

5.2 PROPULSION SYSTEM ANALYSIS (OPTION 2) ..........................................................................................175

6.0 TECHNICAL SHORTCOMINGS AND PORTFOLIO GAP ENGINEERING PLAN FOR 

CONTINUED CONFIGURATION DEVELOPMENT .........................................................................200

7.0 SUMMARY AND CONCLUSIONS ..................................................................................................203

7.1 CONFIGURATION DELIVERABLES ..........................................................................................................203

7.2 REFERENCE VEHICLE REQUIREMENTS ..................................................................................................203

7.3 DESIGN SPACE EXPLORATION AND TRADES ..........................................................................................203

7.4 DOWN-SELECTED CONFIGURATIONS AND SONIC BOOM CONSIDERATIONS .............................................203

7.5 ASSESSMENT OF THE -076E AND DEVELOPMENT/ TECHNOLOGY PORTFOLIO GAPS ................................205

8.0 REFERENCES ....................................................................................................................................207

APPENDIX 1 - DESCRIPTION OF THE 072B .....................................................................................208

APPENDIX 2 - THE BOEING IN-HOUSE MULTIDISCIPLINARY DESIGN AND ANALYSIS 

TOOL .........................................................................................................................................................213

APPENDIX 3 - GIT AND R-R NON-PROPRIETARY TECHNOLOGY ROADMAPPING ...........228
6 
7 

Submitted to NASA LaRC: 

Lori Ozoroski, Technical Monitor 

Peter Coen, NASA Principal Investigator 

Program Management: 

Toni Antani BR&T Peter Radloff BCA 

Principal Investigator: Bob Welge 

Co-Investigators: 

John Bonet Todd Magee Chet Nelson 

Report Authors: 

 Eric Adamson Steve Baughcum John Bonet Robert T. Britt 

 Daniel Chen Steve Hollowell Aaron Kutzmann Todd Magee 

 Greg Miller Alan Mortlock Chet Nelson Josh Stengle 

 Jimmy Tai, GaTech Bob Welge 

Technical contributors in addition to authors: 

Acoustics: 

Kevin Elmer 

Aerodynamics: 

 Alicia Bidwell Dave Bruns Peter Camacho 

 Andy Grothues Brandon Huelman David Hyde 

 Geojoe Kuruvila Eric Unger 

Configuration: 

Kevin Lutke Greg Oakes Tom Smith 

Propulsion: 

Boeing

Edward Lawson Kurt Acheson Wesley Moore Kimberly P. Nguyen 

Georgia Institute of Technology

Russell Denney Christopher Perullo Jonathan Sands 

Pratt & Whitney

Joseph B. Staubach Naushir Bala Peter Robinson Doug Stetson 

Rolls-Royce Liberty Works

Joshua Clough Andrew Barlow Jerry Brines Robert Duge 

Structures/Materials: 

Tom Tsotsis 

Weights: 

Tony Gonzales Paul Miller Brandin Northrop 
8 

0.0 Abstract 

Supersonic airplanes for two generations into the future (N+2, 2020-2025 EIS) were designed: the 100 

passenger 765-072B, and the 30 passenger 765-076E. Both achieve a trans-Atlantic range of about 

4000nm. The larger 765-072B meets fuel burn and emissions goals forecast for the 2025 time-frame, and 

the smaller 765-076E improves the boom and confidence in utilization that accompanies lower seat count. 

The boom level of both airplanes was reduced until balanced with performance. The final configuration 

product is two “realistic”, non-proprietary future airplane designs, described in sufficient detail for 

subsequent multi-disciplinary design and optimization, with emphasis on the smaller 765-076E because of 

its lower boom characteristics. In addition IGES CAD files of the OML lofts of the two example 

configurations, a non-proprietary parametric engine model, and a first-cycle Finite Element Model are also 

provided for use in future multi-disciplinary analysis, optimization, and technology evaluation studies. 

 
9 

0.1 Executive Summary 

This report summarizes the work, the findings, and the contract deliverables by Boeing to NASA LaRC 

under NRA Subtask A.4.7.1 of Contract NNL08AA16B, Order NNL08AA35T.

. 

A. Fundamental Aeronautics Research Program. 

A.4 Supersonics. 

A.4.7 Systems Integration, Assessment and Validation. 

A.4.7.1 Supersonic Concept Development and Systems Studies. 

Work under the contract helped reveal some of the key system-level integration challenges for supersonic 

cruise commercial aircraft. It involved many variations of several innovative and highly integrated 

configurations that applied advanced technologies appropriate for the N+2 years 2020 to 2025 (Entry Into 

Service target). Two of the most promising configurations; a larger 100 passenger, high efficiency airplane 

and a smaller 30 passenger, low boom airplane, were selected for more detailed assessment and design. 

From those, the smaller low boom airplane was chosen to provide a vehicle for an even more thorough 

assessment and more detailed design of a non-proprietary, small, supersonic airliner for delivery to NASA. 

It is accompanied by a substantive description of a viable airplane configuration for exercising alternative 

design and analysis methods. 

Figure 0.1.1 outlines the overall process followed in the course of arriving at the final configuration. The 

first step was arriving at an achievable set of reference vehicle requirements that could lead to an 

economically and environmentally viable supersonic airliner in the specified time frame. That was 

followed by sizing various airframe and engine combinations to meet those requirements. Subsequent 

analysis and design of many other configuration and technology options, both by individual expert and by 

automated Multi-Disciplinary Analysis (MDA) and Multi-Disciplinary Optimization (MDO), investigated 

several different aspects of integrated airframe, engine, payload and mission parameters. Once a final 

configuration was selected, CFD design was used to benchmark the currently achievable cruise 

performance relative to that projected based on Linear Theory design, and to explore whether achieving 

those levels demanded significant modification. A FEM structural sizing and one iteration of re-design was 

performed to help guide future structural design and identify structures related development and technology 

gaps. A higher fidelity MDA; including a highly effective, response-surface based, and non-proprietary 

engine design capability from the Georgia Institute of Technology; was employed in an integrated study to 

balance engine noise at takeoff with overall performance. A complete description, analysis results and 

geometric definition of the final airplane is provided to NASA within this report and separate files. 

The optimized 100 and 30 passenger airplanes and their many relatives that were part of this study provide 

evidence that the achievement of very low boom levels and maximizing fuel mileage efficiency are 

mutually exclusive objectives for optimized airplanes of the N+2 generation. However, hypothetically, the 

two representative configurations chosen could be equally viable N+2 concepts, provided that “express” 

supersonic over land service with higher ticket surcharges provided sufficient economic and social value to 

the airlines and customers relative to a larger airliner that only operates at full speed over water. This is 

especially true if the impact of the higher fuel burn per seat of the smaller vehicle could be offset by 

improvements in low-NOx combustor technology and the use of carbon-neutral bio-fuels. 

Other key results of the study are that designs resulting from Linear Theory, at least those to which Linear 

Theory readily applies, are reasonably good predictors of the performance achievable after design by CFD 

helps to realize the details. Also it is believed that linear-based solutions can be successfully used to 

establish the low-boom potential “goal” of most configurations. Analysis of the jet velocity characteristics 

of the final -076E concept with an optimized engine cycle indicated that it should be possible for such a 

configuration to achieve the minimum community noise standard of 15dB cumulative reduction relative to 

“Stage 3/Chapter 3”. The -15 EPNdB cumulative delta could be achieved using a thrust derate scheme 

without resorting to complex mixer-ejector nozzles like those developed under the HSR program in the 
10 

1990’s. If the required noise stringency moves toward 18-20dB cumulative reduction, a low noise inlet 

concept and some type of low-impact suppressor nozzle are likely required technologies. Finally, 

improvements to boom will demand, to an even greater extent than improvements to basic performance, 

aggressive structures, materials, and controls technology. 

Figure 0.1.1 Roadmap to accomplish the goals and requirements of the NRA task A.4.7.1. 

 

Requirements

• Mission

Goals

• Environmental

• Efficiency

1st Order Analysis

Mission Performance, Boom, airport noise and Emissions

1st 6 Months 2nd 6 Months

Revise Existing MDA Framework for Small SS Aircraft Application

1st Year 

Small SS Aircraft

Selected Concept

Parametric Geometry

MDO Framework

Documentation 

NASA Analysis Ready

MDO Ready

1st Year 

Small SS Aircraft

Selected Concept

Parametric Geometry

MDO Framework

Documentation 

NASA Analysis Ready

MDO Ready MDO Dev

SFW ¶ 2.5.2 NRA 

Typical Example

TOGW

Boom

2020

Preferred Concept

Refined analysis

CFD, Rapid FEM, Boom

Build Parametric 

Geometry Generator

Propul. 

Options

Config. 

Options

Chose a concept

Reqmts & Year 2

Goals

Initial Concept Final Concept
11 

1.0 Market Study and Goals 

The present N+2 study is to represent supersonic airliner concepts that, hypothetically, if market and 

regulatory conditions were favorable, could enter passenger service as early as the year 2020-2025. A part 

of the study is to forecast these conditions, and to construct the requirements and objectives that will guide 

the airplane design. 

1.1 Timing 

Developmental and certification flow-times for recent new airplane programs suggest that roughly 5 to 8 

years of dedicated research and development would be required to achieve appropriate technology levels. 

Accordingly, the several Supersonic Business Jet (SSBJ) development programs that have been appearing 

in the press for several years represent N+1 level technologies, while the last of the large Mach 2.4 High 

Speed Civil Transports of the NASA HSR program of the 1990’s would represent technologies somewhere 

between the NASA N+2 and N+3 levels. The Entry Into Service (EIS) date of the present N+2 falls 

between the EIS assumed for these other existing studies. 

1.2 General Market & Regulation Forecast 

The viability of any future supersonic airplane will depend on its ability to deliver sufficient economic 

value to its passengers, operators and manufacturers while meeting safety, environmental, and operational 

requirements. These will be accompanied by a wide variety of subjective judgments about capabilities and 

requirements (e.g. a 100% consensus on acceptable boom level might never be achieved, despite any 

official ruling), and the essential value of speed. 

During the lifetime of the Concorde, market demand for speed was apparent. Due to premium fares, the 

two operating airlines were able to operate profitably on selected routes in the end, although the direct 

economic returns to its manufacturers were limited. Since the retirement of Concorde in 2003, there are no 

longer any large differences between airliner speeds, and there are few differences between services offered 

to high-yield travelers. But, the disruption in air travel following September 11, 2001 was followed by 

increasing demand from travelers willing to pay a premium price to shorten travel time by flying business 

jets, and an overall trend of increasing delivery of new business jets, as shown in Figure 1.2.1. It was more 

evidence of a continued demand for improved service and speed. The present economic downturn has 

curtailed overall travel and business travel, but projections are for a gradual return to the long-term, prerecession trends, with some variation due to shifts in the energy and investment sectors. That trend 

includes increasing numbers of travelers willing to pay premium fares to save time. Recent air travel 

statistics also indicate that premium fares are paid by about 10% of travelers on scheduled international 

airlines, and Figure 1.2.2 shows a trend of slow, continuous growth in overall premium ticket volume. At 

some point in the future, a critical mass of high value, time-sensitive travelers would likely be sufficient to 

support at least a limited civil supersonic fleet in at least some markets. Initially, this projected demand for 

faster travel might be satisfied by supersonic business jets. Eventually, if the trend continues, supersonic 

airliners of 30 to 100+ passengers may become viable. But an important lesson from Concorde is that, 

even if market demand in 2020-2025 were theoretically sufficient to support a fleet of supersonic airliners, 

directly replacing the Concorde without incorporating significant advancements in fuel efficiency, noise 

and reliability would clearly be a business disaster, and environmentally irresponsible. 

Some compromise in cabin comfort for supersonic speed would likely be economically sustainable because 

flight times would be shorter. The highest density and lowest acceptable level of comfort would probably 

be the present-day coach class interior of a medium range single aisle airplane, but a typical interior of a 

supersonic airplane would likely be similar to the business class configurations of subsonic airliners already 

familiar to the anticipated community of travelers. 
12 

An economic goal for the large HSCT airplanes was to fly profitably with a ticket surcharge of less than 

20% above premium subsonic prices for supersonic airplanes introduced 10 years in the future, and 10% 

for airplanes introduced 20 years in the future. N+2 airplanes have neither the development time, nor the 

efficiency per seat , due mainly to their smaller size and seating capacity, so those are probably not realistic 

goals for N+2. But they are representative of the strong pressure to keep operating and ownership costs 

low; costs that will inevitably be compared to the best subsonic aircraft of the same generation. 

Another important lesson from Concorde, and from subsonic 1st class and other premium fare service—

particularly the all 1st or all business class cabin arrangements--is that matching demand with supply is 

crucial for achieving favorable economics. The number of markets that could support supersonic service, 

the size of a viable supersonic fleet, and the number of seats within a supersonic airliner would all be 

limited. Airplanes between 30 and 100 passengers are assumed for N+2; larger than an SSBJ to improve 

the fuel mileage per seat, but not as large as the 250 to 300 passenger HSR/HSCT airplanes that might be 

more difficult to fill regularly. 

Environmental acceptability will likely be the most important requirement to meet, as it was for the 

HSCT/HSR program of the 1990’s, i.e. 

- No significant impact on stratospheric ozone. 

- Compliance with then-year airport and community noise standards. 

- No environmental damage from sonic booms. 

- Compliance with all then-year requirements for low and high altitude aircraft engine emissions 

(including airport NOx & CO2). 

At both high and low altitude, lower emissions are historically achieved through basic aerodynamic and 

propulsion efficiency and lower weight, and those are good for basic performance as well. Improved 

combustor design will have the largest effect on reducing NOx. But, the susceptibility of the stratospheric 

ozone to NOx released at high altitude introduces cruise altitude as another factor that must be considered. 

Recent studies by Baughcum (Ref 1) and Wuebbles (Ref 2) show that the impact on ozone increases with 

increasing altitude, and that the impact varies linearly with total NOx emitted. These results update similar 

studies during the HSR Program (Ref 3 and Ref 4), taking advantage of recent advances in atmospheric 

knowledge and modeling capabilities. One significant result in the IPCC aviation assessment (Ref 4) was 

that the water vapor emissions from the HSCT (Mach 2.4) could accumulate in the stratosphere and be 

more important than the HSCT's CO2 emissions in terms of climate impact. It is likely that emissions 

would dictate a lower cruise altitude than was assumed during the HSR Program, and most likely below the 

altitude that would minimize fuel burn alone. A balance must be struck between the total emissions and the 

susceptibility of the atmosphere. Lower cruise altitudes could also favor lower cruise Mach numbers, 

easier emergency descent certification, and engine cycle parameters that better match requirements for 

reduced takeoff noise. 

Airframe and engine noise must be addressed through a balanced combination of engine and configuration 

features. Light weight, low wing loading, high lift-to-drag ratio (L/D) at low speed, and high bypass ratio 

engines are important characteristics that should help noise. But pursuing them exclusively could adversely 

affect cruise efficiency, so proper, cycled trades are necessary to realize a balanced compromise between 

features favorable for noise and their impact on cruise performance. And other details, like clean 

arrangements of high lift devices to reduce airframe noise, might also come with a weight or cruise drag 

penalty, and must be assessed before assuming them. During HSCT, the airport noise requirements for 10 

years in the future were Stage III -1 EPNdB sideline, -5 cutback and -1 approach. At 20 years, the forecast 

was for -4 to -6 sideline, -8 to -10 cutback, and -5 to -6 approach. For N+2 the noise goal was set to an 

overall -15 EPNdB, with -3 sideline, -7 cutback and -5 approach. Figure 1.2.3 points to an increasing 

number of airports setting increasingly stringent noise restrictions beyond the FAR 36 requirements, so 

noise appears to be a subject of growing importance and awareness to regulators and the population. 

Some progress in engine noise suppression technology has been demonstrated, as shown in Figure 1.2.4, 

but forecasts from the HSCT/HSR Program for the present day have not been met. A part of the shortfall 

can be attributed to a focus on much larger bypass ratio subsonic engines, and minimal investment in 

supersonic technologies. It exposes an area of risk where development is required. 
13 

It is critical to achieve an “acceptable” level of sonic boom, but this will almost certainly come at some cost 

in efficiency of the airplane. Concepts for boom reduction and boom shaping generally demand slender, 

lightly loaded, and uniquely tailored airframes that are not as efficient structurally or aerodynamically as an 

unconstrained design. Cruise efficiency already argues for relatively thin lifting surfaces and slender 

bodies with challenging Aero- Propulsion- Servo- Elastic (APSE) characteristics, and low boom concepts 

add to that challenge. It will demand material, structural mode control, and flight stability & control 

capabilities beyond those already assumed for keeping weight and drag low on an airplane optimized for 

performance. Avoiding boom by selectively flying at or below “threshold Mach” sacrifices some of the 

speed advantage, and would somewhat compromise the cruise efficiency if the airplane were to be designed 

for a wide range of cruise speeds. Aggressive strategies for boom are constrained by their impact on 

performance, and on the risk that technical solutions would not be available in the N+2 time frame. 

The advantages of twin engine configurations for all but the largest subsonic airplanes--reduced acquisition 

and maintenance costs, fuel efficiency and structural efficiency of the installation--are assumed to apply to 

supersonic configurations as well. Twin engines are assumed for the N+2 airplanes, setting these airplanes 

apart from Concorde and the larger 4 engine configurations of HSR/HSCT. 

Seating density and weight of passenger accommodations in most 1st class cabins today result in about 3 

times the fuel per seat as coach class due to inefficiencies of reduced scale, and typical business jets burn 

roughly twice the fuel per passenger as airliners. This is illustrated in Figure 1.2.5. So premium fare 

service in a supersonic airplane could be offered without extremely tight interior arrangements while still 

comparing favorably with subsonic economics from the operator’s perspective. But public scrutiny of fuel 

consumed by air travel has been growing, and will likely continue to grow, so a supersonic N+2 airplane 

should avoid being a major outlier in fuel use or associated emissions. An achievable N+2 target of about 

20% reduction in fuel burn per passenger per nautical mile over current business jets and 33% reduction 

from Concorde (and even better in more typical mixed-Mach operation) is identified in Figure 1.2.5. 

Similar fuel burn data versus passenger count in Figure 1.2.6 depicts how the 100 passenger N+2 

supersonic fuel burn target was chosen using a curve fit to be 0.29 lb fuel per passenger mile. Figure 1.2.7 

demonstrates that the resulting range varies significantly, depending on the payload and configuration. 

Trans Atlantic range is assumed essential for capturing high-density markets that support premium fares 

and involving time-sensitive travel, and for showcasing the advantage of speed. 4000nmi will be the range 

required for the N+2 airplanes. 

We assume that some fuel--petroleum, bio, other synthetic or some blend--with roughly equivalent heating 

value to present jet fuel will be available throughout the operational lifetime of an N+2 airplane. But fuel 

cost will continue to rise, emphasizing fuel efficiency as an essential part of keeping the airplanes 

economically viable. 

Safety will not be compromised, and certain lessons since the original HSR/HSCT have raised the 

awareness of designers and regulators to issues that must be addressed during the design. A sample is 

included in Table 1.2.1, showing how the N+2 airplanes will require additional features and design effort. 

An initial step to an operational supersonic airliner might necessarily be a SSBJ that paves the technical and 

regulatory path to a safe, economical, and certifiable airplane. A smaller predecessor would not only 

provide lessons, but also set in place some regulations, and, more importantly, help establish the framework 

within which regulations would be set. 

 
14 

1.3 Initial Market/Design Requirements & Objectives 

The full list of Market/Design Requirements & Objectives (M/DR&O) for a 30 passenger airplane is 

attached in Table 1.3.1, and for a 100 passenger airplane in Table 1.3.2. Some key points for are the 

following: 

- 1.6-1.8 cruise Mach. Low enough to enhance engine durability, BPR for takeoff & approach 

noise, restricted cruise altitude, ordinary temperature structural materials. High enough to offer 

significantly shorter flight times. 

- 55,000ft cruise altitude, or lower, for emissions & simpler cabin certification. 

- 0.95 or lower cruise Mach below 39,000ft for ATC margins. 

- No supersonic speeds below 41,000ft for ATC margins. 

- Supersonic fuel burn less than 0.3 lb/pax/nmi, for economics and emissions on the 100 

passenger airplane. 

- Sonic boom as low a practical, less than Concorde over water, and possibly “boomless” flight at 

“threshold Mach” or lower over land. 

 
15 

Table 1.2.1 Areas of safety and certification that have received additional attention in recent years. 

 
16 

Table 1.3.1 Reference Market/Design Requirements and Objectives for 30 pass. low boom airplane. 

Description Item Reqirement Objective Comment

1 EIS EIS 2025 2020

2 Type Cert Type Cert FAR Pt 25

3 Op Cert Op Cert FAR Pt 135

4 Cruise Speed Cruise Mach 1.6 1.8

Initial Subsonic Cruise 

Altitude Capability (M=.95) ISCAC 39,000 ft 50,000 ft

Initial Supersonic Accel 

Altitude Capability ISCAAC 41,000 ft 45,000 ft

5

Initial Maximum Cruise 

Mach Altitude Capability ICAC 45,000 ft 50,000 ft

6 Maximum Altitude MAC 55,000 ft 53,000 ft

emissions & 

decompression

7 Time and Distance to Climb Time (ISA+ 10 deg C, MTOW) <= 40 min to 45,000 ft <= 25 min to 50,000 ft

8 Payload/Range Range 4000 nm 6000 nmi @ M1.6

Mission Payload 30pax 45pax 210lb/pax + 9,000lb cargo

ETOPS 120 min

9 Takeoff Field Length Balanced TOFL 10000 ft 9000 ft SL/86 deg @ MTOW

10 Landing Field Length 7000 ft 6000 ft SL/86 deg @ MLW

11 Approach Speed Vapp @ MLW 160 kt 150 kt

12 Takeoff Weight MTOW 180,000lb

13 Max Zero Fuel Weight MZFW Basic OEW+ payload

14 Max Landing Weight MLW MZFW + FAR reserves

15 Environmental fuel consumpton best achievable best achievable

proportional to CO2, NOx 

& water vapor emissions

Steady over land boom in corridor 0.8psf 0.5psf

Steady over land boom unrestricted 0.35psf 0.25psf

Steady over water boom 2psf 1.5psf

Community Noise Ch 4 minus 10 EPNdB Ch 4 minus 14 EPNdB

Ramp Noise <= 82 dBA @ service locations <= 80 dBA

Interior Noise <= 90 dBA <= 80 dBA

16 Longitudinal CG Design CG Limits Available aero limits

17 Flight Operations Balance Pasenger seating limitations None

Fwd CG limit MZFW CG + fwd OEW variation

Aft CG Limit OEW CG + aft OEW variation

18

Ground Operations 

Balance Static: Tip up unlimited passenge movement

19 Taxi Ride Comfort Ride Smoothness No worse than 737NG

20 Turning Radius Turning Radius ICAO Class B ICAO Class C

21 Airframe Design Life Airframe Design Life 20K hrs/10,000 cycles/20 yrs 30K hrs/15,000 cycles/25 yrs

22 Cabin Passengers 30, dual class bus. + econ.

Width 2-3 AB bus. class

Seat Bottom 18" 20" bus / 18" econ Data from Payloads

Aisle 16" 21" bus / 19" econ

Seat Pitch 40" bus / 30" econ 40" bus / 32" econ

Galley Carts 4

Cargo load 9,000 lb

Cargo volume 900ft^3

Pressure Altitude 6000 ft

Flight Crew 2 2

Cabin Attendant Per regs

23 Engines Number 4

Min Supercruise TBO 2500 hrs 3500 hrs

Failures Meet FAA Adv Circ 10 128A

24 Fuel System Fuel type ASTM 1655 Jet A, A-1, JP-4, E7

Active CG management Fuel transfer system

25 Landing Gear Runway loading ICAO Class B

26 Door Sill Height Door Sill Height Between 10ft & 20ft

27 Auxiliary Power APU 120 kVA Scaled from 787

28 Electrical Power Electrical power required 300 kVA VFG Scaled from 787

29 Flight Deck Pilot training reqd Similar to conventional airliner

30 Flight Controls Control philosophy FBW

31 Gear Breakaway

Prevent damage to primary 

structure and breech of fuel 

tanks.

32 Foreign Object Injestion No worse than 737NG.

33 Fuel Inerting Required.

34 Engine containment

Critical systems and structure 

within burst zone must be 

redundant and separated.

35

Control Authority with 

Failed systems Single failures will be surviable.

36 Passenger Cabin egress

Door number & proximity to meet 

FAR Part 25
17 

Table 1.3.2 Reference Market/Design Requirements and Objectives for 100 passenger high efficiency 

airplane. 

Description Item Reqirement Objective Comment

1 EIS EIS 2025 2020

2 Type Cert Type Cert FAR Pt 25

3 Op Cert Op Cert FAR Pt 135

4 Cruise Speed Cruise Mach 1.6 2.0

Initial Subsonic Cruise 

Altitude Capability (M=.95) ISCAC 39,000 ft 50,000 ft

Initial Supersonic Accel 

Altitude Capability ISCAAC 41,000 ft 45,000 ft

5

Initial Maximum Cruise 

Mach Altitude Capability ICAC 45,000 ft 50,000 ft

6 Maximum Altitude MAC 55,000 ft 53,000 ft

emissions &

decompression

7 Time and Distance to Climb Time (ISA+ 10 deg C, MTOW) <= 40 min to 45,000 ft <= 25 min to 50,000 ft

8 Payload/Range Range 4000 nm 6000 nmi @ M1.6

Mission Payload 100pax 100pax 210lb/pax + 9,000lb cargo

ETOPS 120 min

9 Takeoff Field Length Balanced TOFL 10000 ft 9000 ft SL/86 deg @ MTOW

10 Landing Field Length 7000 ft 6000 ft SL/86 deg @ MLW

11 Approach Speed Vapp @ MLW 160 kt 150 kt

12 Takeoff Weight MTOW 300,000lb

13 Max Zero Fuel Weight MZFW Basic OEW+ payload

14 Max Landing Weight MLW MZFW + FAR reserves

15 Environmental fuel consumpton <= 0.35lb/nmi/pax <= 0.30 lb/nmi/pax

proportional to CO2, NOx 

& water vapor emissions

Steady over land boom in corridor 0.8psf 0.5psf

Steady over land boom unrestricted 0.35psf 0.25psf

Steady over water boom 2psf 1.5psf

Community Noise Ch 4 minus 10 EPNdB Ch 4 minus 14 EPNdB

Ramp Noise <= 82 dBA @ service locations <= 80 dBA

Interior Noise <= 90 dBA <= 80 dBA

16 Longitudinal CG Design CG Limits Available aero limits

17 Flight Operations Balance Pasenger seating limitations None

Fwd CG limit MZFW CG + fwd OEW variation

Aft CG Limit OEW CG + aft OEW variation

18

Ground Operations 

Balance Static: Tip up unlimited passenge movement

19 Taxi Ride Comfort Ride Smoothness No worse than 737NG

20 Turning Radius Turning Radius ICAO Class B ICAO Class C

21 Airframe Design Life Airframe Design Life 20K hrs/10,000 cycles/20 yrs 30K hrs/15,000 cycles/25 yrs

22 Cabin Passengers 100, dual class bus. + econ.

Width 4 AB bus. class

Seat Bottom 18" 20" bus / 18" econ Data from Payloads

Aisle 16" 21" bus / 19" econ

Seat Pitch 40" bus / 30" econ 40" bus / 32" econ

Galley Carts 4

Cargo load 15,000 lb

Cargo volume 1,500ft^3

Pressure Altitude 6000 ft

Flight Crew 2 2

Cabin Attendant Per regs

23 Engines Number 4

Min Supercruise TBO 2500 hrs 3500 hrs

Failures Meet FAA Adv Circ 10 128A

24 Fuel System Fuel type ASTM 1655 Jet A, A-1, JP-4, E7

Active CG management Fuel transfer system

25 Landing Gear Runway loading ICAO Class B

26 Door Sill Height Door Sill Height Between 10ft & 20ft

27 Auxiliary Power APU 200 kVA Scaled from 787

28 Electrical Power Electrical power required 500 kVA VFG Scaled from 787

29 Flight Deck Pilot training reqd Similar to conventional airliner

30 Flight Controls Control philosophy FBW

31 Gear Breakaway

Prevent damage to primary 

structure and breech of fuel 

tanks.

32 Foreign Object Injestion No worse than 737NG.

33 Fuel Inerting Required.

34 Engine containment

Critical systems and structure 

within burst zone must be 

redundant and separated.

35

Control Authority with 

Failed systems Single failures will be surviable.

36 Passenger Cabin egress

Door number & proximity to 

meet FAR Part 25
18 

Figure 1.2.1 Trends in new business jet deliveries. 

Figure 1.2.2. Trend in premium ticket volume. 
19 

Figure 1.2.3 Increasing numbers of airports setting noise restrictions more stringent than FAR36. NAP 

means Noise Abatement Procedures are in effect. Airport data compiled by Boeing Noise Engineering, 

Community Noise. 
20 

Figure 1.2.4 HSCT noise suppression, demonstrated (solid black symbols) and projected. 

Figure 1.2.5 Fuel burn per passenger for several types of airplanes. 
21 

Figure 1.2.6 Fuel burn versus passengers for various types of airplanes. 

Figure 1.2.7 Fuel burn versus range for various types of airplanes. Different and unspecified payload 

accounts for much of the wide variation in range. 

 
22 

2.0 Concept Development 

Overview 

Results of the Market Study served in all subsequent concept development activity to set requirements and 

help evaluate trades. Their influence will appear directly or implicitly throughout this document. 

The overall scale of a 100 passenger N+2 airplane falls between the larger HSR/HSCT and smaller SSBJ, 

as shown in Figure 2.0.1, with no overlap with either. Little direct comparison to previous work will be 

possible without relying on basic scaling, and few of the detailed configuration analyses (e.g. addressing 

APSE, planform optimization, engine-airframe matching, etc.) from the other projects will be directly 

applicable. As a result, the N+2 work will necessarily include additional studies, and areas of risk that 

might have already been covered in the other programs. Figure 2.0.2 provides a notional view of how 

aggressive payload, range and boom targets have the potential to leave no feasible territory. Requirements 

and goals will likely need to evolve in order to produce an airplane for N+2. 

Some enabling technologies have matured significantly since HSCT/HSR, and are now on production 

airplanes. For example, the 787 has an “all composite” structure, nozzle chevrons and improved inlet noise 

treatment. Active maneuver/gust load alleviation systems, flight deck technology (including 

EVS/SVS/XVS), adaptable FBW/mixed function control surfaces, and the efficiency of turbo-machinery 

have all advanced. And designers enjoy improvements to CFD, FEM and optimization tools; better 

aerodynamic testing methods and instrumentation; and improved aeroelastics tools and methods. But 

development of most technologies specific to supersonics has slowed significantly or stagnated since the 

end of HSR, and in some cases HSCT ideas and lessons have been lost. SSBJ and related R&D have 

carried a few technologies forward, but, as a sign of things to come, many old lessens had to be re-learned. 

In general, the final HSCT configurations are actually 4-7 years advanced from the N+2 time-frame. 

The first steps in the present work toward an N+2 airplane began with a “level 0” analysis to establish a 

notional “green” Concorde, with about 75% of the TOGW while carrying 50 first class or 100 dual class 

passengers. The engine was selected from a matrix of available study engines by installing them on scaled 

versions of two of the best final HSR configurations at constant wing loading and thrust loading. A 

preliminary configuration was assembled from promising features and sized. A 30 passenger candidate 

was also developed, starting with existing Boeing and NASA trade-study results, to provide an alternative 

with lower boom. Both the large and small configurations were developed further using a combination of 

interactive and automated design methods. 

2.1 Initial Concepts 

The initial sizing study began with two concepts derived from airplanes of the NASA HSR/HSCT Program. 

One was the 765-070A (figure 2.1.1) based on the 1080-2015 configuration, or “2015-TC”. The other was 

the 765-071B (figure 2.1.2) based on the 1080-2154 configuration, or “2154 HISCAT” that promised lower 

noise and increased structural stiffness near the engines. The red dashed lines in Figure 2.1.1 suggest an aft 

deck that might provide a similar increase in stiffness for the 765-070A, and will appear subsequently on 

the 765-072B discussed below. 

The Boeing performance sizing code ASAP was used to size thrust and area at 1.8 Mach with 47 

passengers and fixed 300,000lb TOGW to meet the fuel burn target of 0.3 lb fuel per passenger-mile on a 

long range mission. The sizing was repeated for several Boeing proprietary study engines, and it was the 

Boeing Study Engine 2, or “BSE-2”, engine that proved most favorable. Figures 2.1.3 and 2.1.4 are 

representative sizing charts showing the results for one of the engines. These sizing results led to 

configurations that were essentially 76% scale versions of the originals. 
23 

Compared to the HSCT engines, these BSE-2 engines have higher bypass ratio and lower fan pressure ratio 

appropriate for lower cruise Mach. That should improve both subsonic and supersonic SFC, and takeoff & 

approach noise should also improve. With lower nozzle pressure ratio and jet velocity, nozzles are 

expected to be lighter, simpler, and less expensive to maintain. But being twin engine airplanes, these N+2 

concepts will demand higher thrust-to-weight ratio for engine-out performance than the quad HSR/HSCT 

airplanes. Regulatory restrictions for traffic and boom would also argue for higher thrust-to-weight, despite 

any engine differences, because less efficient profiles of speed vs. altitude are expected for N+2 airplanes 

than were assumed during HSR/HSCT. Higher bypass ratio and thrust-to-weight both mean larger 

nacelles, as shown in figure 2.1.5, so drag and weight—including some for the likely additional aft deck 

area to stiffen the engine support structure--will offset some of the gain. And while the inlet flow path 

feeding a single engine will be simpler, the inlet will be longer for acoustic treatment to address increased 

fan noise, for variable geometry for takeoff and transonic airflow control, and for FOD shielding. 

Varying passenger count (i.e. changing the weight of passengers and their accommodation in proportion) 

and cruise Mach, Figure 2.1.6 shows the range that the 765-071B could fly if TOGW were fixed at 

300,000lb. Cruising at high subsonic Mach number for part of a mission could extend range, providing 

some compensation for the lost speed. But choosing to slow to Mach 1.2 would cost both speed and range, 

indicating that the advantage of possibly flying at “threshold Mach” to avoid boom must be weighed 

against the penalty in performance. The value of speed has not been formulated, so it is not clear whether 

that would be a favorable trade. Figure 2.1.6 also illustrates how range and passenger count trade. Clearly 

there is a cost in range for hauling more passengers and less fuel, but seat-mile economics generally favor 

more passengers, so the best design will just meet the required range. 

The result of sizing the wing area, thrust and fuel load of the 765-071B with 100 passengers at Mach 1.8 

for various different ranges is shown in Figures 2.1.7. It demonstrates that an N+2 configuration can 

achieve the CO2 emissions goal of 0.9 lb/pax/nmi and the fuel consumption goal of 0.30 lb/pax/nmi with a 

300,000lb airplane flying a 4000nmi mission. Similar sizing was performed assuming 25, 50 and 75 

passengers, and demonstrates that the per-passenger economics clearly favor the largest passenger counts. 

Arguing for lower (or not any higher than 100) passengers are the practical lessons from the Concorde and 

ordinary first class service--that an airline might have difficulty keeping load factors high on flights with a 

premium price. Figure 2.1.7 also includes the result for flying each of those sized airplanes at other Mach 

numbers, and it reiterates the message of Figure 2.1.6, that flying at high subsonic speed improves 

performance, but that the low supersonic speed of 1.2 Mach is detrimental. 

Compared to the 1080-2154 HISCAT at Mach 2.4, the 765-071B had 13% lower (worse) cruise L/D and 

17% lower (better) SFC. The main contributors to lower L/D were the lower Reynolds number and 

increased relative excrescence that comes with a smaller airplane, larger nacelles to house larger engines, 

longer inlets, lower aero technology projections assuming fewer opportunities on the N+2 configurations 

and less optimism regarding stability augmentation and structural mode control systems. The gain in SFC 

is primarily from the reduced Mach and higher bypass, and with that comes a favorable reduction in noise. 

A more complete description of the 765-070A is in figures 2.1.8 through 2.1.12, and of the 765-071B in 

figures 2.1.12 through 2.1.16. Comparing some of the data listed in figures 2.1.9 and 2.1.14, the 765-071B 

has a slightly higher cruise L/D of 8.73 and lower fuel burn of 0.280lb/pax/nmi compared to to 8.59 and 

0.288lb/pax/nmi for the 765-070A, so it is a slightly more efficient airplane. 

Interior arrangements in Figures 2.1.11 and 2.1.16, configured by Boeing Payloads Engineering, 

demonstrate that dual class arrangements of 100 seats are feasible within each cabin. The 765-070A has a 

longer and narrower single aisle arrangement that loads closer to the overall airplane CG. The wider and 

shorter cabin of the 765-071B just allows dual aisles, affording some structural advantage from this and 

from eliminating a pair of doors, but with some disadvantage of narrowing rapidly near the structural joint 

to the wing box. 

The weights summaries in Figures 2.1.10 and 2.1.15 illustrate that some structural weight in the wing and 

fuselage swap between the two configurations (i.e. the 765-070A has a smaller and lighter wing but 

heavier body), but the 765-071B has 2,850lb greater overall structural weight, primarily because of the 
24 

larger and heavier wing that is not completely compensated by the lighter body. Some of that difference 

would disappear if a more prominent aft deck were added to the 765-070A to stiffen it comparably. The 

765-071B has 1,600lb less weight associated with passenger accommodations (partly from eliminating the 

pair of doors), and is less than 1% heavier in overall OEW. 

Figures 2.1.12 and 2.1.17 tell that the boom of the 765-071B at 103.72 PLdB is better than the 765-070A at 

109.16 PLdB, but with little margin to meet future over-water regulation. Neither airplane would likely be 

allowed to fly near population at cruise Mach. So, while these two airplanes meet the cruise performance 

and emissions goals, they do not meet the goal for boom. To pursue boom, two paths were followed. One 

was to configure a 100 passenger derivative of these two airplanes that incorporates their favorable features 

with an emphasis on promising avenues for reducing boom. That became the 765-072B that will be 

described in section 2.2. The other approach was to configure a smaller airplane that sacrifices some of the 

payload and fuel efficiency, but would allow significant boom reduction. That became the 30 passenger 

765-076E described in section 2.3. 

2.2 100 Passenger High Performance Transport 

After demonstrating with the 765-070A and 765-071B concept airplanes that the performance targets for 

N+2 were feasible, the 765-072 was assembled using them as a guide, but with some emphasis on features 

that would assist in integrating an airplane with lower or “softened” boom. It was initially constructed and 

analyzed interactively within the Boeing MTA conceptual design tool, and several cycles of interactive 

design led to the 765-072B configuration. A python-based MDA was also employed using brute force runs 

of thousands of cases, single-objective MDO, and multi-objective (Pareto) optimization. Some of these 

studies and their results will follow. Analysis and design at higher fidelity was accomplished with the 

ModelCenter-based MDA and MDO. 

2.2.1 Initial Trade Studies 

Initial trade studies employed a combination of interactive and automated design. The Boeing Multidisciplinary Trade and Analysis (MTA) tool provided the interactive capability. It is an Excel interface to 

configuration definition and Level 0-1 analysis tools. It enables ad hoc programming to enforce the 

relationships unique to a study (e.g. a particular parametrization of a planform); and a push-button interface 

(inputs, execution & outputs) to the analysis tools ANLZ, NFWD, SKFR, T080, DragJ, MDboom/MDplot, 

Zephyrus/Loudboom, a SEEB boom worksheet, QWICKO, and a mission performance simulator. Typical 

studies involved varying either geometry or the F-function, and letting the drag, boom PLdB, or closeness 

to target (e.g. target F-function when varying geometry, or target geometry when varying F-function) 

guide next steps. Success depended critically on the creativity and insight of the designer, but allowed 

broad flexibility in configuration options and design methods. 

Automated study and MDO of the initial concepts were with a multi-disciplinary analysis tool built as a 

python-scripted interface to the same aerodynamics, drag and boom analysis tools. It allowed one-off runs, 

brute-force sampling of design space, or optimization under the direction of Design Explorer. It could be 

driven either through parametrized geometry, or through parametrized F-function. Typical objectives were 

L/D (at single Mach or weighted sum), boom PLdB (overall, forward, or aft), and overall body volume--a 

surrogate for volume of the passenger cabin, baggage, cockpit, some fuel, systems, and gear wells. While 

not as flexible as the interactive MTA tool because it demands an explicit statement of any new 

configuration or design approach, this automated tool facilitated numerical optimization and massive 

sampling of the parametrized design space. 

Promising candidates from either design approach were fitted with an interior arrangement by Boeing 

Payloads Engineering according to industry standards for airliners and regional jets. These airplanes also 

received a more thorough weights and performance assessment, including an attempt to re-size the wing 
25 

and engine using ASAP. And each was cycled through both the interactive and the automated tools as a 

safeguard against spurious results unique to one of the methods. 

Higher fidelity analysis and design was accomplished using an MDA assembled in ModelCenter. A 

parametric geometry generator fed a series of the same linearized aerodynamics tools, followed by NPSS, 

CART3D, CASES, and QWICKO. The usage of NPSS pre-dated the full capability employed 

subsequently on the 765-076E, but took advantage of some of the preliminary design-space surveys and 

response surface information about the engines produced by Georgia Tech. 

Early work was focused on selecting appropriate objectives and design variables for subsequent 

optimizations. Figure 2.2.1.1 shows an example, where the Mach number used within T080 for constrained 

minimum wave drag was either fixed at the cruise Mach number of 1.8 or free to be any value between 1.1 

and 2.0. The initially surprising result that design Mach numbers other than cruise could lead to better L/D 

illuminates the unsurprising conclusion that minimizing volume wave drag does not minimize overall drag 

(volume wave + lift induced wave + skin friction + interference). The design Mach within T080 became a 

variable in subsequent optimizations. 

Another preliminary study explored the effect of considering L/D evaluated at the transonic thrust pinch 

(typically Mach 1.1) in the L/D objective. Studies of the 765-072A using ASAP revealed that L/D at the 

thrust pinch had about 50% greater impact on the cycled airplane than L/D at cruise. The effect that Mach 

at which to evaluate L/D had on optimization was demonstrated using 2-objective Pareto optimizations, 

with one objective to maximize the body volume, and the other to alternatively maximize the L/D at cruise, 

the L/D at pinch and a weighted sum of 40% cruise and 60% pinch. Results in Figure 2.2.1.2 show that 

maximizing L/D at cruise alone leads to a substantially lower L/D at pinch than if L/D at pinch alone were 

maximized, and vice a verse. It also reveals that including the L/D at pinch produces the same result 

whether the L/D at cruise is included or not. Subsequent optimizations used the weighted sum. And 

whenever planform trades included span, a third component, the L/D at subsonic Mach 0.95, was added to 

the weighted sum to ensure that it would influence the objective. 

Optimum body cross-section for the fixed wing planform of the 765-072A was investigated using a 2-

objective Pareto optimization to maximize L/D and overall body volume using body variables, resulting in 

an estimate of the best L/D available for any given body volume. Data of this type was essential for 

configuration trades involving the body because, lacking a carefully arranged interior for each candidate, it 

was not yet obvious which would hold the required 100 passengers. And no more than 100, since excess 

volume would clearly impact L/D, as demonstrated in both Figures 2.2.1.2 and 2.2.1.3. But achieving the 

highest L/D depended, in part, on how much area ruling could be accomplished in the aft fuselage. 

Consequently, configurations on the Pareto frontier had unacceptably narrow bodies at the wing spar, 

where the critical structural joint between wing and fuselage occurs. So optimization was repeated with the 

body constrained to be at least some minimum depth at a location near the wing spar, and that location was 

varied to produce the family of curves in Figure 2.2.1.3. They show the cost in L/D for increasing the 

fuselage depth near the spar. Figure 2.2.1.4 shows examples of the resulting optimal area distributions for 

fixed overall body volume of 14,000 cubic feet, revealing that the aft fuselage depth could be doubled 

(green vs. red) at a cost of about 1% in L/D. The corresponding structural advantage is at least a weight 

savings, but potentially salvages an otherwise infeasible structural arrangement. It also moves the 

passenger load aft, closer to the overall CG. Results from this study helped establish the overall volume of 

the body and aft body shape on the 765-072B. 

A study of the outboard wing sweep and span revealed that supersonic L/D at this level of analysis is not 

very sensitive to planform, and that subsonic L/D is predictably sensitive to span through induced drag. A 

more comprehensive figure of merit that credits the subsonic segments of the mission would respond to 

these variables, but static and dynamic structural and control issues argue against any changes in sweep or 

span toward improving L/D. The results were the same whether simply varying sweep and span (i.e. 

nothing else varied except for the T080 minimum wave drag fuselage design through fixed constraint 

points), or optimizing and letting all the body variables free. 
26 

The effect of sweep and span on boom PLdB was also both modest and accompanied by adverse 

consequences for the structure. Results for a simple parameter variation (i.e. without optimization) are 

shown in Figure 2.2.1.5, depicting a range of variation in boom PLdB no greater than what could be 

achieved by modest variation of the body cross-section. It is likely that simultaneously designing the 

planform and body to minimize boom would result in practically the same level, and exploiting any small 

advantage offered at this level of analysis by a particular span and sweep would probably require higher 

fidelity tools. 

The effect of altitude and Mach number on boom PLdB is shown in Figure 2.2.1.6 for the 765-072A 

cruising at 270,000lb. Unfortunately, boom PLdB appears to be maximized at the chosen cruise Mach of 

1.8, although it is not clear that any general conclusion can be drawn from this particular case. The sudden 

change in the trends near 55,000ft is explained in Figure 2.2.1.7, where the double shocks in the front 

coalesce before reaching the ground. It is not surprising that cruise altitude can affect the ground signature 

this way, but it demonstrates one of the many challenges to achieving desired boom characteristics. 

A direct application of the multi-objective optimizer was to use 3-objective Pareto optimization to trade 

L/D, body volume and boom. L/D was typically either the weighted sum of 40% cruise plus 60% pinch, or 

32% cruise plus 48% pinch plus 20% subsonic. Figure 2.2.1.8 shows two dimensions—volume and L/D—

from such a 3-objective optimization, and included on it are results from the corresponding 2-objective 

optimization in those same two dimensions. The agreement between the two results at the volume-L/D 

frontier is excellent, lending confidence in 3-objective results. In Figure 2.2.1.9 all three dimensions are 

portrayed, along with contours of constant PLdB that help illustrate a steep trade in PLdB near the volumeL/D frontier. And scatter in the PLdB values for the 2-objective results is more evidence that only a small 

price in L/D or volume would have to be paid to lower boom PLdB measurably. Away from the volumeL/D frontier, the trends are more gradual but consistent--PLdB gets worse when L/D or volume increase. 

These trends are made clear in figures 2.2.1.10a & b, showing orthogonal views of the three-dimensional 

data, with iso-curves of the parameter normal to each plot. 

An allied dataset optimizing the same 3 objectives, but with more design degrees of freedom at the nose 

and aft end of the body, is shown in Figure 2.2.1.11. As expected, the additional freedom leads to better 

values of the objectives, but the interesting result shown here is evidence that the aft part of the boom is the 

dominant contributor to overall PLdB. The forward and aft components of the boom PLdB were estimated 

starting with the whole ground pressure signal that MDboom predicts, then discarding half of it, and then 

analyzing in MDplot in the usual way, as if it were the whole ground pressure signal. Either a falling or 

rising hyperbolic tangent step of the form (1-tanh(x))/2 was used to clip the forward or aft parts of the 

pressure signal smoothly to minimize any spurious content. The centers of the tanh() steps were aligned 

with the centroid of pressure fluctuations, and the width (i.e. reciprocal of the scale of the argument) was 

set to 10ft. This forward-aft decomposition of the boom PLdB provides some opportunity for the optimizer 

to discover and exploit some direct relationships between geometry and boom. In some later studies when 

one end of the boom was consistently dominant (typically the aft end, but not always), minimizing PLdB 

for that end alone was an objective. Just as in Figures 2.2.1.8 through 2.2.1.10, these data also indicate that 

overall PLdB gets worse when either L/D or volume increase, although plotting all the data together in 

Figure 2.2.1.11 masks these individual trends. 

At the conclusion of these and several cycles of interactive design using the Boeing MTA tool, the 765-

072B configuration emerged, shown in Figure 2.2.1.12. Its fuel efficiency was promising, and additional 

design and analysis was performed at higher fidelity using the ModelCenter-based MDA/O, as described 

below. An important general result from these and several other similar optimization studies on 100 

passenger airplanes was that the design space afforded little opportunity to improve boom without 

sacrificing performance and demanding development beyond the N+2 level. That result was reinforced 

with the manual development of 765-078A & B, using the MTA tool, described in Section 2.2.10. 
27 

2.2.2 Trade Space Studies Using MDAO Model of the 072B Concept 

To provide a comprehensive analysis of the 765-072B, and to gain confidence in the decision to accept it as 

the 100 passenger candidate, a higher fidelity analysis and engine matching effort was performed. It 

involved a more detailed series of system level trades on 072 and 070 type concepts. In order to capture the 

interactions between the various analysis disciplines (geometry, aero, weights, propulsion, mission 

performance, takeoff performance), an in-house MDAO model was used. (This MDAO model is described 

in Appendix 2). 

Two MDAO models for the 072B and 070 (2015-TC) type aircraft were developed. The following trade 

approach was followed: 

• Species trade using both MDAO models 

• System level trades using the best species MDAO model 

o Number of passengers 

o Cruise Mach 

o Cruise Range 

o Engine cycle (from 6 preselected cycles) 

For all the trades, a supersonic non-stop mission profile was used to determine the performance of the 

concepts. This mission profile, shown in Figure 2.2.2.1A and 1B, is parameterized to allow for variations 

in cruise range and Mach number. 

For the species trade, the 765-070 type and 765-072B type species were compared and the best species was 

selected for the remaining trades. 

For the engine cycle trades, a series of six candidate engines cycles, three from P&W and three from RR, 

were evaluated. These were provided to the MDAO model in the form of installed thrust and fuel flow 

tables for a reference thrust. In deciding how to choose the engine cycles for the system level assessment, 

the design team believed that jet noise and engine diameter would be two competing characteristics. Jet 

noise was expected to be the dominant noise component, so jet velocity (Vj) served as a surrogate for 

propulsion noise. It was generally understood that we wanted a low diameter engine for best supersonic 

aerodynamic performance and we wanted a low Vj to minimize the jet noise. However, because of the 

fundamental physics of turbine engines, minimizing Vj results in a large diameter engine, and vice versa. 

Therefore, Georgia Tech prepared three engine cycles for each of the engine company’s proposed 

technology level representing a range of diameter and Vj. These engines are designated 

• Low Diameter #1 and #2, 

• Low Vj #1 and #2, and 

• Compromise #1 and #2 

for the purposes of the trade studies. Trade studies selecting the Compromise Engine # 1 for the other 

trades discussed below is given in Section 2.2.8. 

For each of the trade study points, the concept was “optimized” to maximize the Figure of Merit (FOM) of 

PAX*nmi/lbs of fuel by varying the wing area and engine size (reference thrust). When the wing area is 

scaled, the MDAO model recalculates all the aerodynamics, weights, and resulting performance. When the 

reference thrust is scaled, the MDAO model changes engine size/weight, along with the thrust and fuel 

flow tables. When the engine size changes the nacelle will also change. So the aerodynamics must be 

recalculated. All of these will affect the performance. The optimum aircraft must meet the following 

constraints. 

• Greater than 300 fpm rate of climb (ROC) anywhere in the mission 

• Less than 10,000 ft balanced field length (with engine failure) 

• Sufficient fuel volume in the wing (desired) or wing and fuselage (required) 

• Less than 78 psf wing loading (W/S) 

The wing loading constraint was selected based on historical data which indicated that keeping it below 78 

psf would result in an approach speed below 155 kts at typical maximum landing weight, and Initial Cruise 

Altitude Capability (ICAC) unrestricted by buffet boundary at subsonic and supersonic Mach. The results 

of a two variable optimization can be depicted graphically with all the constraints in a carpet plot or thumb 
28 

print. Graphical results are easy to comprehend and give insight into what constraints are driving the 

optimum design and what alternatives the design space allows. 

Following the trades, a preferred concept was selected and a fairly simple multi-disciplinary optimization 

(MDO) was conducted to demonstrate the parametric capability of the MDAO model. At the time the trade 

studies were conducted, the aerodynamics, propulsion, and performance analyses in the MDAO model 

were still being validated. This exercise was completed after the trades and the performance of the 

preferred concept received minor updates. The changes to the model as a result of the validation did not 

affect the relative trade study results or the conclusions. 

2.2.3 Species Trade 

The 765-070 type and 765-072B type species were compared in the species trade. For the purposes of 

performance assessment, the primary difference in these two concepts will be in the cruise aerodynamics. 

Figure 2.2.3.1 shows a comparison of the lift over drag (L/D) ratios for these two concepts. The -070 has 

approximately a 5% lower maximum L/D and the differences increase slightly with increasing lift 

coefficient (CL). The aerodynamic differences are caused primarily by the horizontal tail on the -070, as 

well as small differences in wing sweep, thickness, and aspect ratio. 

In a separate analysis, the Boeing team determined that the airport noise requirements (10-20 EPNdB 

below Stage 3) had a high probability of being met without a noise suppressor on the nozzle if the Vj could 

be kept at or below 1100 fps. During the prescreening performance runs, we observed that none of the six 

candidate engines could achieve 1100 fps Vj at maximum takeoff power. We also observed that, in 

general, the engines were sized to get through the transonic pinch point without diving the aircraft or using 

afterburners. In doing this, they had excess thrust for takeoff. This meant that the engines could be 

throttled back (or derated) for takeoff as a means to reduce the Vj. Figure 2.2.3.2 depicts how this works. 

For this example, an aircraft with the engine sized for transonic ROC > 300 fpm could achieve a balanced 

field length of ~ 5500 ft using 100% power (blue line). At 100% power the engine would have a Vj of 

1530 fps (magenta line). By reducing the power to 60%, the same aircraft achieves a 10,000 ft balanced 

field length. At 60% power the Vj is approximately 1200 fps. However, for engine-out performance, the 

FAR allows the thrust to be increased by 10% and does not require that the noise requirements be met. 

This means that for meeting the noise requirements we can use the jet velocity for an all-engine takeoff at 

10% lower thrust. The resulting Vj is 1100 fps, meeting the goal for the study. Simple constant derate is 

assumed in this study, recognizing that Takeoff Field Length or noise could be reduced through a fully 

optimized thrust lapse starting with full thrust a brake release, as in the “PLR” system during HSCT. 

Figure 2.2.3.3 takes this concept a step further. In this example, the engine sized for transonic ROC, meets 

the 10,000 ft balanced field distance at 77% power. At that power setting, even accounting for the 10% 

FAR thrust increase, it will not meet the 1100 fps Vj goal. So the thrust must be increased (blue dotted 

line) until it does meet the 1100 fps goal. This will occur when the engine has enough thrust to meet the 

10,000 ft balanced field length at 66% power. The carpet plots shown for the trades will be using a 10,000 

ft balanced field distance constraint for a partial power takeoff which meets the 1100 fps Vj goal. 

Figure 2.2.3.4 shows the carpet plot for the 100 passenger 072B species with Compromise #1 engine. Each 

of the points in this plot meet a Mach 1.6 cruise and a 4000 nm range. The objective is to maximize FOM. 

The carpet is plotting thrust from 45,000 to 80,000 lbs and wing area from 3000 to 5000 ft2

. The blue line 

shows the 300 fpm ROC constraint. The red line shows that meeting the 10,000 ft balanced field distance 

with the engine de-rated (throttled) to meet 1100 fps Vj is more constraining than ROC. The orange line 

shows the 78 psf W/S constraint and the green line shows the wing fuel volume constraint. The crosshatched side of the constraint lines shows the infeasible portion of the design space. The point where the 

balanced field distance and W/S constraints cross represents the maximum FOM concept that meets all of 

the constraints. This optimum concept has a FOM of 3.7, which is better than the project goal of 3. It has a 

wing area of 4600 ft2

 and a thrust of 55,000 lbs. If we were to relax the derated balanced field constraint 

then the optimum concept would move up the the intersection of the ROC and W/S constraints and the 

FOM would increase to 3.8. 
29 

Figure 2.2.3.5 shows a similar carpet plot for the 100 passenger -070 species with Compromise #1 engine 

meeting a Mach 1.6 cruise and 4000 nm range. Thrust and wing area were selected to vary the same as the 

-072B and the FOM was plotted at the same scale as it was for the -072B to depict the fact that the design 

space for the -070 species collapses down to a lower FOM than the -072B species. As before, the optimum 

design which meets all the constraints occurs at the intersection of the balanced field distance and W/S 

constraints. This concept has an FOM of 3.36, a wing area of 4550 ft2 and a thrust of 59,500 lbs. The 

increased thrust required was due to the higher drag depicted in Figure 2.2.3.1 shown previously. 

The species trade results are summarized in Figure 2.2.3.6. This figure shows that the -072B species has 

higher FOM, lower takeoff gross weight (TOGW) and lower thrust per engine. 

The conclusion of this trade is that the -072B species is the preferred species for the rest of the trades. 

2.2.4 Number of Passengers Trade 

The -072B species was laid out with a 100 passenger configuration. The passenger trade additionally 

considered concepts with 50 and 25 passengers. For this trade, the Compromise #1 engine was used. 

Initially the cruise Mach was set at 1.6 and the range was set at 4000 nm. Later, the combined trade will 

look at varying Mach and range, along with the number of passengers. In order to conduct the passenger 

trade, 25 and 50 passenger fuselages had to be manually laid out. The fuselage cross-section constraints for 

two (blue), three (red), and four (gold) abreast seating is shown in Figure 2.2.4.1. These cross-sections 

were combined into seating cylinders for new 25 and 50 passenger fuselages, as shown in Figure 2.2.4.2. 

The goal was to keep a constant overall fineness ratio and the resulting overall lengths are shown in the 

figure. The fuselage constraints for each of these three passenger layouts were coded into the MDAO 

model, so that the user merely had to select 25, 50, or 100 passengers and the model would automatically 

reconfigure the fuselage. 

Figure 2.2.3.4 from the previous species trade shows the carpet plot for the 100 passenger 072B, having an 

FOM of 3.7. Figure 2.2.4.3 and 4 show the carpet plots for the 50 and 25 passenger 072B concepts, 

respectively. On the 50 passenger concept the constraints for ROC and balanced field distance lie almost 

on top of each other, but the ROC is slightly more constraining. So the optimum concept occurs at the 

intersection of the W/S and ROC constraints. This concept has an FOM of 2.3. On the 25 passenger 

concept, the balanced field distance and the ROC constraints cross at a wing area of approximately 3300 

ft2

. At the intersection of the W/S constraint, the balanced field distance is more constraining. The green 

line in Figure 2.2.4.4 shows that the 25 passenger concept does not have sufficient wing volume for fuel. 

Therefore, some of the fuel is placed in the fuselage (below the passenger compartment). The orange line 

shows the limit for the combination of wing and fuselage fuel volume. The resulting FOM for 25 

passenger concept is 1.3. Figure 2.2.4.5 shows a summary of the passenger trade results. As expected, the 

FOM significantly improves with increasing passenger count. The 100 passenger concept has almost three 

times the FOM of the 25 passenger concept. Since the larger number of passengers results in a larger 

aircraft, the TOGW and thrust per engine also increase with the number of passengers. 

2.2.5 Cruise Mach Trade 

The Mach trade will use the -072B concept with the Compromise #1 engine, initially the 100 passenger 

concept with a 4000 nm range will be compared. Later the Mach number will be combined with variations 

in number of passengers and range. Figure 2.2.3.4 from the species trade shows the carpet plot for the 

Mach 1.6 072B, having an FOM of 3.7. Figures 2.2.5.1 and 2 show carpet plots for Mach 1.8 and 2.0, 

respectively. For Mach 1.8 the intersection of the balanced field distance and wing loading constraints 

defines the optimum aircraft with an FOM of 3.55. The Mach 2.0 concept is constrained by the intersection 

of the W/S and ROC constraints with an FOM of 2.55. The carpet plot in this figure was flipped for clarity 

in displaying the constraints. Figure 2.2.5.3 shows a summary of the Mach trade results. There is very 
30 

little difference (~4%) between the Mach 1.6 and 1.8 cruise points. On the other hand, cruising at Mach 2 

significantly degrades the FOM (~30%) versus Mach 1.6. Since reduced travel time is a principal motive 

behind these supersonic airliner studies, there are other figures of merit, such as productivity (block time) 

or passenger preference that would favor selecting the Mach 1.8 cruise speed. While this particular engine 

cycle and airframe shape shows the Mach 1.6 to have the best FOM, there may be other engine/airframe 

combinations that will give a slight edge to Mach 1.8. 

2.2.6 Range Trade 

The Range trade will use the -072B concept with the Compromise #1 engine, initially the 100 passenger 

concept with a Mach 1.6 cruise will be compared. Later the range will be combined with variations in 

number of passengers and Mach number. Figure 2.2.3.4 shown previously from the species trade shows 

the carpet plot for the 4000 nm 072B, having an FOM of 3.7. Figures 2.2.6.1 and 2 show carpet plots for 

5000 nm and 6000 nm, respectively. The optimum concept for the 5000 nm range occurs at the 

intersection of the W/S and balanced field constraints, with an FOM of 2.95. The 6000 nm concept is not 

really feasible. None of the points meet either the wing loading or derated balanced field constraints. In an 

effort to find a feasible solution, additional points were run at larger thrust and wing area. Various analyses 

failed to converge at these excessively large wing areas and thrust, so it was not possible to find a feasible 

solution without changing the configuration or planform, or sacrificing other parameters. If the W/S and 

balanced field constraints are ignored, the resulting FOM is 2.55. Figure 2.2.6.3 shows a summary of the 

range trade. Only the 4000 nm range results in an FOM above the project goal of 3. 

2.2.7 Combined Range, Mach, # Passengers Trade 

Figure 2.2.7.1 shows a 27 point matrix (3 X 3 X 3) of the Range, # passengers, and Mach trades. To 

visualize the results, the FOM is plotted versus range and a separate curve is shown for each of 25, 50, and 

100 passengers. Then a separate plot is shown for each cruise Mach (1.6, 1.8, 2.0). Starting at a cruise 

Mach of 2.0, none of the combinations of range and # passengers will reach the FOM goal of 3. Further, 

6000 nm range is not feasible and the 5000 nm 100 passenger point does not meet the W/S constraint. 

Next, we see that the Mach 1.6 and 1.8 plots are very similar. None of the 6000 nm points meet the W/S 

constraint. Only one range/passenger combination exceeds the project FOM goal of 3. It is 4000 nm range 

and 100 passengers. So ignoring the boom for now, the recommended combination would be 4000 nm 

range, 100 passenger and either Mach 1.6 or 1.8 cruise. 

2.2.8 Engine Cycle Trade, Fixed Engine Decks 

The engine cycle trades were all run using a 100 passenger -072B species with a Mach 1.6 cruise and 4000 

nm range. Six engine cycles were run (see Section 3). The six cycles and their corresponding carpet plots 

figures are shown below. 

• Low Vj #1: Figure 2.2.8.1 

• Low Vj #2: Figure 2.2.8.2 

• Compromise #1: Figure 2.2.8.3 

• Compromise #2: Figure 2.2.8.4 

• Low Diameter #1: Figure 2.2.8.5 

• Low Diameter #1: Figure 2.2.8.6 

For the Low Vj #1 engine, the optimum point occurs at the intersection of the W/S and ROC constraints 

with a FOM of 4. The balanced field distance constraint does not show on the plot because all the points 

meet the constraint with a derate power of 79%. This engine starts out with the lowest Vj at 100% power, 

so we would expect the minimum amount of derate required to meet the 1100 fps Vj goal. For low Vj #2 

engine, the optimum point which meets all the constraints occurs at the intersection of the balanced field 

and W/S constraints. It has an FOM of 3.21. The required derate power is 66% in order to meet the 1100 

fps Vj. There is quite a bit of difference between the balanced field constraint and the ROC constraint for 
31 

this engine. If the balanced field constraint is ignored, then the FOM rises to 3.57. For Compromise #1 

engine, the optimum point occurs at the intersection of the balanced field and W/S constraints. It has an 

FOM of 3.7. This is the same airframe/engine combination that was used for the Mach, range and 

passenger trades. The required derate power is 60% in order to meet the 1100 fps Vj. For the Compromise 

#2 engine, none of the points were able to meet the 10,000 ft balanced field distance with the required 54% 

derate power. Ignoring this constraint, the FOM would be 3.8. For the Low Diameter #1 engine, none of 

the points were able to meet the 10,000 ft balanced field distance with the required derate power. Ignoring 

the balanced field constraint, the FOM is 3.9. Finally, for the Low Diameter #2 engine, again none of the 

points were able to meet the 10,000 ft balanced field distance. Ignoring the balanced field constraint, the 

resulting FOM is 3.5. 

Figure 2.2.8.7 summarizes the results of the engine cycle trade. The blue bars are for a non-derated takeoff 

and the number in the blue bar shows the Vj achieved. The maroon bars are for the derated takeoffs. Only 

three of the six cycles were able to achieve a 10,000 ft balanced field distance with the engine derated to an 

1100 fps Vj. Based on FOM alone, the Low Vj #1 engine cycle would be the best, followed by the 

Compromise #1 and Low Vj #2. These three cycles all exceed the project FOM goal of 3 and meet all the 

constraints. The Compromise #1 engine, however, has the lowest thrust per engine. It was also the opinion 

of the team members that the Low Vj #1 engine cycle would be more difficult to integrate because of the 

large fan diameter. Therefore, Compromise #1 is the recommended engine cycle from this study. 

Subsequent improvements and errors fixed in the MDAO analysis changed the predicted performance 

level, but not the conclusions. 

2.2.9 Trade Study Summary and Recommendations 

Following these studies using the MDAO, improvements and corrections were made to the tool, changing 

the level of the answers (see Figure 2.2.9.1), but not the conclusions. The trade study results show that the 

-072B species is preferred over the 070 species based on FOM. Only the combination of 100 passengers 

and 4000 nm range were able to achieve the project FOM goal of 3. Either Mach 1.6 or 1.8 are acceptable 

for cruise and there is very little difference between their FOM, but with the (as yet un-quantified) value of 

speed arguing for faster. Finally, the Compromise # 1 cycle was selected for the purposes of producing a 

computer aided design (CAD) model of the 072B airplane, shown in Figure 2.2.9.2 

2.2.10 Boom Reduction 

The hopeful low boom successor to the two initial 100 passenger concept aircraft was to be the 765-072. 

But after several manual iterations and MDO studies, it became clear that the configuration afforded little 

opportunity to lower the boom significantly without substantially compromising the weight or drag, and 

adding development risk beyond the N+2 level. As a rough order of magnitude estimate, the goal was a 

100 passenger airplane that has 35% lower fuel burn than Concorde, meets “Stage 5” noise criteria, and has 

less than 83 PLdB boom. Compared to the Concorde, it would need lower wing loading, greater span, a 

compound-sweep planform (possibly with variable sweep), moderate BPR engines with suppressor nozzles, 

very low NOx engines at low weight with commercial durability, 25% reduction in TOGW at 400nmi more 

range, 25-35% empty weight reduction, and a length of about 300ft. It requires light-weight, stiff and 

probably actively controlled structure with an effective structural l/d that exceeds that of any flying 

airplane, or even any HSR/HSCT designs, and represents technology beyond the N+2 time frame. 

In light of the challenges, an attempt was made to soften the boom of the 765-072B using the 765-078 

configurations. Both the A and B variants achieved lower boom through softening, but at the expense of 

performance, as they both exceeded the fuel burn requirement. 

The 765-078 preserved many of the essential characteristics of the 765-072B. The overall length increased 

to 224ft, and some details of the body area distribution were free to change, but, as shown in Figure 

2.2.10.1, some constraints were enforced to preserve the viability of the 100 passenger interior. One was to 
32 

preserve the width of the cockpit. Another was to ensure that the forward cabin and aft body crosssectional areas remained large enough to hold a galley/lav, or 2-abreast seats, or APU and vertical tail tiein; and that at least 14,200ft^3 total volume was preserved. A cruciform tail was added in an attempt to 

alter the aft lift signature, and the aspect ratio and sweep of the canard were raised to help drag. Wing 

planform was varied, but area was not allowed to change significantly. 

The planform, lift distribution and body shape were designed using the Boeing MTA tool. The first step 

involved approximating the F-function of the 765-072B with a simplified model, then varying this simpler 

F-function iteratively to design a softened boom. The resulting F-function then became the target, and a 

first approximation to the airplane geometry that could produce it was calculated using the Abel transform. 

From that geometry began a series of geometric design iterations; where the resulting F-function, boom 

signature, PLdB and drag were monitored. The focus was on boom softening, with no particular attention 

paid to wave cancellation to lower the boom. Often the target F-function was re-examined and sometimes 

re-designed in light of results. 

After several design cycles, two candidates with fairly good boom PLdB emerged, the 765-078A and 765-

078B. Figure 2.2.10.2 shows the 765-078A in 3-view and its principal characteristics. Favorable boom 

reduction was achieved by creating a double shock in the forward part of the ground signature, and some 

subtle improvement in the aft, shown in Figure 2.2.10.3. MDboom and Zephyrus both report about the 

same boom level of 105 PLdB. That was clearly an improvement to the 109+ PLdB of the 765-072B, but it 

came at a cost of 11% higher fuel burn and 260nmi less range. 

One more attempt to improve the boom was through softening the ground signature. The configuration was 

the 765-078B, and it is essentially the same at the 765-078A at the scale of the figure. Apparently minor 

modifications resulted in a significantly smoother front shock, and clipped aft shock on the ground, as 

shown in Figure 2.2.10.4. An even lower boom of less than 98 PLdB was achieved, but the penalty grew to 

17% higher fuel burn and 410nmi range shortfall compared to the 765-072B. 

The 765-078 A and B provided evidence that a penalty in performance can be expected for reducing the 

boom of an efficient cruise airplane. With the 765-072B just meeting the performance requirements, it 

appears that it can afford no further boom reduction. 

2.3 Thirty (30) Passenger, Lower Boom Transport 

The 765-076 model was introduced to study a smaller airplane with potential for substantially lower boom 

than the 100 passenger 765-072B, but accepting a loss in MTOGW and payload capacity, and with worse 

fuel burn and emissions per seat. It was initially sized to have a TOGW of 180,000lb with a similar wing 

loading and thrust loading as the 765-072B. 

The E variant of the 765-076 is shown in figure 2.3.1. Overwing nacelles are to help achieve the lift and 

volume distributions required for low boom, but with some penalty in wave drag due to lift. Other benefits 

are some shielding of fan noise, and a significantly lower risk from runway, slush or tire FOD. Canard 

surfaces were omitted to avoid ingesting their vortices. V-tails assist with integrating the overwing 

nacelles, and theoretically add some (as yet unproven) virtual length to the lift distribution by their height 

above the wing, assuming a substantially aft CG for boom. 

2.3.1 Trade Studies 

The 765-076 matured into the E model through primarily manual iteration through the MTA tool. The 

approach began with analytical and optimization studies to arrive at a target F-function that produces 

reasonably good boom PLdB, then using the Abel transform to get a first approximation to the geometry 

that might produce it. Then the study proceeded much like the study of the 765-078, where geometry was 

varied toward matching the resulting F-function to the target. The target evolved according to the insight 
33 

gained by numerous boom analyses into the practical limits of using airplane geometry to control details of 

the realized F-function. 

An early study of the 765-076 concept was an attempt to gage the potential for boom reduction while 

paying little attention to the penalty in performance. It began with an aft wing configuration scaled to 

180,000lb MTOW. The python-based MDA was used with Design Explorer to minimize PLdB by varying 

the body cross-section distribution, wing span and outboard wing sweep. Figure 2.3.1.1 shows results from 

that study. L/D and overall body volume were constrained not to fall below values not far from the seed 

configuration. In figure a) the benefit of allowing overall length to vary is demonstrated, but, in this study 

the length was not independent of volume because the cabin cross-section was also roughly constrained to 

preserve space for constant seating capacity. Thus, longer airplanes had more total volume, so their boom 

suffered, despite the otherwise favorable effect of length on boom. When wing was fixed (red +) a small 

improvement was possible. But when the wing was free to change span and sweep, significantly lower 

boom was achieved. While initially promising, all of the configurations with better boom also had 

significantly higher span or sweep or both, assuring a penalty in weight (that was not assessed) and 

exaggerating the APSE challenges. 

Another path of study was to drive a parametrized, constrained F-function with the optimizer to establish 

some off-body pressure targets that achieve good boom by various means (e.g. reduced overpressure, boom 

shaping, multiple shocks), as assessed by MDboom. A closed-form expression for the equivalent area 

distribution, derived in familiar form for a piece-wise linear F-function using the Abel transform, estimates 

the physical geometry and lift that could produce a similar off-body pressure distribution. Figure 2.3.1.2 

shows a typical parametrized F-function. All points were free to vary in station and amplitude, but 

amplitudes were scaled collectively within each of the four groups shown to solve for the following four 

constraints on the equivalent, transformed geometry: 

Overall lift must match weight, setting cross-section at aft end. 

The slope of the aft end must match that of a typical supersonic airplane. 

The maximum of the equivalent cross-section (lift plus volume), as fraction of aft cross-section. 

Positive geometric volume at the aft end of lift. 

Figure 2.3.1.3 makes those constraints in the transformed geometry apparent. Other constraints (e.g. 

overall equivalent volume or the station of the area maximum) and various combinations were also possible 

to satisfy explicitly. 

Several optimizations were performed using Design explorer to drive toward minimum boom PLdB using 

this parametrized F-function formulation. In the cases that will be shown, the points defining the Ffunction were divided into three sets, and constraints on lift, slope of the equivalent geometry at the aft end, 

and overall volume of the equivalent body were met explicitly. The seed to each successive optimization 

was a promising point (typically the optimum) from a previous result. Bounds on the variables were 

successively adjusted to free any variables that had been up against them at the prior optimum. The data 

are shown in Figure 2.3.1.4. Figure 2.3.1.4a shows how PLdB was minimized by minimizing volume. It is 

not a surprising result, and re-iterates that basic parameters like volume (and lift, length) are the principal 

drivers of boom, so design of other details becomes something like an exercise in avoiding penalties from 

poor execution. The lower edge of the data spanning the plot represents an approximate Pareto frontier, 

providing an estimate of the minimum boom for any given volume. Figure 2.3.1.4b shows the forward and 

aft parts of the boom analyzed separately, and demonstrates that the forward part dominates. Two of the 

optimizations used a subset of the variables so that little would change in the aft boom, while attention 

would be focused on the forward boom. 

The F-function of the two points with the lowest boom levels are shown in Figure 2.3.1.5. One surprising 

feature of the lowest, 82 PLdB case is the unusually small amplitude to the overpressure at the front. But, 

the other case has a more familiar “spike” in pressure (F-function), so it might not be an important feature. 

This method of driving the F-function provides no assurance that the corresponding pressure distribution 

could be realized in the signature of a real airplane, and the steep signals and large amplitude at the aft end 

of both of the curves in Figure 2.3.1.5 might, therefore, be spurious. Nevertheless, the ground signature, as 
34 

propagated from 50,000ft by MDboom, is plotted in figure 2.3.1.6. Surprisingly, the overpressure at the 

front of the lowest boom is substantially higher than the higher boom. But both have achieved a somewhat 

gradual initial shock, followed by a substantially more gradual rise to the maximum. The favorable aft 

boom levels measured in Figure 2.3.1.3 are evidently the result of fairly mild aft amplitude that relaxes 

essentially without a shock. That behavior is certainly favorable for boom, but probably spurious. 

2.3.2 High Fidelity Analysis and Design. 

Just prior to finally selecting the 765-076E, and while the more slender and lower boom 765-076F was still 

under consideration, TRANAIR design was used to trim the 765-076F, and to measure how much change 

from the Linear Theory design would be required to achieve the predicted performance. Four designs were 

attempted. The first allowed only wing and tail camber and twist to be free. The second added freedom of 

the aft body to the list. The third added the freedom of the nacelle shape to the list. And the final added the 

tail cant angle to the list, encompassing all the degrees of freedom. The resulting configuration from each 

design was also analyzed at Mach 0.95, and the third design at various tail angles to explore the effect of 

CG. Contours of local Mach number are shown in Figure 2.3.2.1 for the baseline before design and after 

the fourth design, the one with all degrees of freedom. With very little observable difference in local Mach 

distribution, an 11 count reduction in drag was achieved, bringing the designed case much closer to the 

predicted drag from Linear Theory. 

The last three rows of Table 2.3.2.1 show the effect of CG on pressure drag, according to Tranair, on the 

third configuration designed by Tranair. Moving the CG aft costs drag for a trimmed airplane, opposite to 

conventional wisdom for ordinary subsonic airplanes, and perhaps adverse for minimizing boom. Carrying 

lift aft is typically good for boom, but apparently that comes at a price in drag. 

To do an integrated performance analysis of the -076E, we needed a new species MDAO model. This 

model had more limited parametric capability than the -072B in order to keep the shaped boom signature 

characteristics. Specifically, the external shape of the vehicle was not allowed to scale. A non-parametric 

GEODUCK geometry model of the -076E was integrated into the same MDAO framework as the -076E 

model. The aero characteristics of the -076E were calculated with the same codes as the -072B. The only 

difference being that, since the geometry of the 076E was fixed, the fuselage volume distribution 

optimization to minimize wave drag was not done. 

A point-of-departure engine cycle was needed for the NPSS non-proprietary engine cycle optimization 

(Section 3), and a proprietary BSE 04B cycle with a reference thrust of 33,000 lbs was used. This engine 

had a fan diameter of 66 inches. To meet the boom signature requirements, the TOGW was required to be 

180,000 lbs or less. This meant that if the fuel required to meet a 4000 nm range caused the TOGW to 

exceed 180,000 lbs, then the range must be reduced until the TOGW was equal to 180,000 lbs. The initial 

performance runs showed that the thrust would have to be increased to 43,000 lbs to make it through the 

transonic pinch point. This level of thrust equates to a fan diameter of 75.4 inches with this BSE 04B 

engine cycle. Even with this extra thrust, the aircraft could not get through the pinch point with a 180,000 

lb TOGW. So the fuel was removed until the aircraft achieved 300 fpm rate of climb capability at the 

transonic pinch point. This resulted in a TOGW of 171,800 lbs and a range of 3200 nm. The fan diameter 

of 75.4 inches was set as the point of departure for the non-proprietary engine cycle optimization in Section 

3. Figure 2.3.2.2 shows a drawing of this configuration, and Figure 2.3.2.3 is the performance summary. 

The 765-076E is described in more detail in Section 4. 

2.3.3 Boom Reduction. 

With boom reduction a goal for the 30 passenger airplane, it was necessary to explore the potential. With 

that in mind, the 765-076F and G were configured with the same wing planform and engines but with 

significantly longer noses than the E. Cabin volume and overall volume were held nearly constant in an 

attempt to preserve the payload. The 765-076F configuration is shown in Figure 2.3.3.1, and the loft 
35 

compared to the 765-076E in shown in Figure 2.3.3.2. In addition to 20ft more of overall body length, 

nearly all of which was added to the nose, two streamlined volumes were added to the tips of the vertical. 

Their purpose is both as a housing and fairing for ballast for stabilizing the fin, and as an unproven 

technology for tailoring the boom. They have potential of either favorable wave cancellation, or of 

effectively lengthening the body by means of their height projected along the standard, forward leaning 

Mach cut for assessing boom. Perhaps the latter effect could only be favorable if the volume of the pods is 

offset by volume saved in the fins, themselves, but that is a possibility if the stabilizing effect of the pods 

were sufficient. 

Figure 2.3.3.3 shows results from MDboom and Zephyrus at Mach 1.8 at 49,000ft and the same 162,000lb 

weight used to assess the 765-076E. 85 PLdB (the result from Zephyrus) might be achievable with the 

765-076F. That is better than the 765-076E at 91 PLdB, but the advantage in boom comes with a penalty 

of 10% in both increased fuel burn and lost range. It also comes with additional APSE risk from the long, 

slender nose, and the additional overall weight cannot be accurately predicted without a more sophisticated 

and careful look at the loads and structural sizing. That weight would increase the penalty on efficiency 

and range, and decrease the benefit to boom. 

Since adding 20ft was good for boom, adding more could be even better. The 765-076G was given an even 

longer nose in an attempt to find what length would be required to reach 70 PLdB. The result was an 

additional 30ft, to an overall length of over 200ft, or more than 30% longer than the 765-076E. Of course, 

all the disadvantages seen in the 765-076F are exaggerated, with about 14% higher fuel burn and less range 

than the 765-076E, and much greater risk of APSE issues; and all before assessing the impact of the certain 

increase in weight. 

While the particular cases of the 765-076F and G might be too few to conclude that boom and drag would 

trade against one another, that conclusion can be drawn from the data in Figure 2.3.3.4. It shows results 

from a 2-dimensional Pareto optimization to minimize boom PLdB and maximize L/D, starting from the 

765-076E. The body design variables are used, and cabin volume and overall body volume are constrained 

to be at least as much as the baseline. The results show that the 765-076E is near the Pareto frontier, and 

that, indeed, boom and drag will trade against one another on the frontier. While these results indicate 

some slack in the baseline, it is probably close enough to the frontier to justify leaving Linear Theory 

design for higher fidelity methods. 
36 

Figure 2.0.1 Size and weight of 100 passenger N+2 studies compared to other supersonic airplane 

studies. 

Figure 2.0.2. Notional TOGW required to meet range versus range for various different boom levels. 
37 

Figure 2.1.1 a. 76% scale “2015-TC”. b. 765-070A external definition. 

Figure 2.1.2 a. 76% “2154 HISCAT”. b. 765-071B external definition. 

a. 76% 2154 OML b. 765-071B OML 

b. 765-070A OML a. 76% 2015 OML 
38 

Figure 2.1.3 Level 0 sizing of 2015-TC. 

1080-2015, 300k, Mach 1.8

RR RB577-260-sdb3

2015 300,000 lb 47 passengers

 NAPD Aero Performance

Tue Apr 29 10:17:25 2008

Engine Thrust - SLST (lb)

Wing Area (sq.ft)

22500

22000

21500

21000

20500

4000 4100 4200 4300 4400 4500 4600 4700 4800 4900 5000

RANGE=5700

5750

5800

5850

5900

5950

6000

6050

6100 6200 6150

6250

6250 6300

6350

6400

6450

6500

6550

6600

6650

6700

6750

6800

DFV=0.000

FSGRAD=0.005

TM=0.000

TOFL=8500

MTOW...........lb 300000 

Sw..........sq.ft 4355 

Sh..........sq.ft 0 

Sv..........sq.ft 0 

SLST Thrust....lb 21179 

Range..........nm 6408.0 

Payload........lb 12222 

SLST/eng.......lb 21179 

BET/eng........lb 25924 

OEW............lb 143767 

Block Fuel.....lb 132429 

Reserve Fuel...lb 13260 

Fuel Vol Req...lb 165090 

Fuel Vol Avail.lb 165089 

Block Time.....hr 7.61 

Climb Time....min 42.19 

ICAC...........ft 53103.00

Thrust ICAC....ft 53103 

Buffet ICAC....ft 99999.00

Fact AE TOFL...ft 8434 

NOTE: 1st and 2nd Seg Grad

 based on 8500 ft FL

1st Seg Grad..... 0.0627 

2nd Seg Grad..... 0.0831 

 Takeoff LE ..deg 30.00 

 Takeoff TE ..deg 12.00 

 Landing LE ..deg 0.00 

 Landing TE ..deg 30.00 

 Approach CL..... 0.7257 

 Vappr.......keas 133.0 

 ASAP Version B3_4 March-18-2008

CT

 CD

Mach Number

0.03

0.02

0.01

 1.0 1.5 2.0 2.5

Climb Rate (ft/min)

Mach Number

8000

4000

 0

 1.0 1.5 2.0 2.5

POT

AVG

CD

CT
39 

Figure 2.1.4 Level 0 sizing for 2154 HISCAT. 

1080-2154 HISCAT, 300k, Mach 1.8

RR RB577-260-SDB3

1080-2154 300,000 lb 47 passengers

 NAPD Aero Performance

Mon May 5 14:15:46 2008

Engine Thrust - SLST (lb)

Wing Area (sq.ft)

23000

22800

22600

22400

22200

22000

21800

21600

21400

21200

21000

3600 3700 3800 3900 4000 4100 4200 4300 4400 4500 4600

RANGE=5550

5600

5650

5700

5750

5800

5850

5900 5950 6000 6200 615061006050 6250

6300

6350

6400

6450

6500

6550

6600

6650

DFV=0.000

CLIMBTIME=60

TM=0.000

TOFL=9000

MTOW...........lb 300000 

Sw..........sq.ft 3844 

Sh..........sq.ft 0 

Sv..........sq.ft 0 

SLST Thrust....lb 21682 

Range..........nm 6325.0 

Payload........lb 12222 

SLST/eng.......lb 21682 

BET/eng........lb 26540 

OEW............lb 145337 

Block Fuel.....lb 130941 

Reserve Fuel...lb 13220 

Fuel Vol Req...lb 163546 

Fuel Vol Avail.lb 163549 

Block Time.....hr 7.68 

Climb Time....min 58.00 

ICAC...........ft 50742.00

Thrust ICAC....ft 50742 

Buffet ICAC....ft 99999.00

Fact AE TOFL...ft 8321 

NOTE: 1st and 2nd Seg Grad

 based on 9000 ft FL

1st Seg Grad..... 0.0735 

2nd Seg Grad..... 0.0928 

 Takeoff LE ..deg 30.00 

 Takeoff TE ..deg 10.00 

 Landing LE ..deg 0.00 

 Landing TE ..deg 30.00 

 Approach CL..... 0.7798 

 Vappr.......keas 137.1 

 ASAP Version B3_4 March-18-2008

CT

 CD

Mach Number

0.03

0.02

0.01

 1.0 1.5 2.0 2.5

Climb Rate (ft/min)

Mach Number

8000

4000

 0

 1.0 1.5 2.0 2.5

POT

AVG

CD

CT
40 

Figure 2.1.5 Engine integration differences between HSR/HSCT quad and 765-070A. Some aft deck 

area (dashed red line) is probably required for both strength and stiffness of the structure supporting 

the engine. 

HSR 

Low-BPR 

QuadJet 

New 

High-BPR 

Twinjet 
41 

Figure 2.1.6 Maximum range of the 765-071B for various cruise Mach and passengers. Thrust and 

wing area sized at constant design weight of 300,000lb. 
42 

a) 

b) 

c) 

Figure 2.1.7. a. CO2, b. fuel burn, and c. MTOW for the 765-071 sized for particular range using 

wing area, thrust, and MTOW, for and 100 passenger payload. 
43 

Table 2.1.8 Characteristics data of 765-070A 

Figure 2.1.8 General arrangement of 765-070A. 

109.82ft 

241 ft

model DATA TABLE (as drawn airplane)

765-070A Wing Horizontal Canard Vertical

ITEM ESDU Total Tail Tail

Area to CL 4035.392 4403.4 323.92 110.90 208.32

Exposed 323.92 110.90 189.04

Reference 4035.392 323.92 110.90 189.04

Aspect Ratio 2.99 3.43 1.199 1.20 2.122

Taper Ratio 0.121812 - 0.283 1.199 0.235

LE Sweep angle 0.00 54.2 54.2 52.61

Dihedral, TE 12 55 0 0 0

T/C 0.024 0.024 0.035 0.000

Tail Volume 0.203 0.138 0.065 0.0891

Span , in 1317.84 903.808 236.49 138.37 169.93

Root Chord, In 786.1323 788.66 307.47 179.90 260.24

Tip Chord, in 95.76 56.000 87.01 50.91 61.23

M.A.C. IN 531.0202 746.643 217.77 127.42 181.27

X 1/4 mac 1713.044 1557.946 2629.15 463.05 2722.82

Y, Zmac 243.4895 80.108981 69.596 99.43

Tail Arm, IN. 916.10 1250.00 1009.7726
44 

Figure 2.1.9 Features of 765-070A. 

 

•Design TOGW= 300,000 lbs 

•Design Cruise Mach=1.8 

•Sref= 4035 sqft 

•Design Range (100-118 pass.)=4000 nmi 

•Cruise L/D=8.59 

•60K lb thrust class “BSE2” engine 

•Fan diam. =84” 

•HSCT-like LE, TE devices 

•3-surface pitch control /trim/ ride qual. 

•Fuel burn 0.288 lb/seat nmi @ 100 pass. 

 0.244 lb/seat nmi @ 118 pass. 

 (=4 seat nmi /lb) 
45 

Table 2.1.10 Weights summary of 765-070A. 

 
46 

Figure 2.1.11 Interior arrangement of 765-070A. 
47 

a)

b)

EQUIVALENT AREA COMPARISON

-10

90

190

290

390

490

590

0 20 40 60 80 100 120 140 160 180 200 220 240

STA ~ X ~ FT

EQUIVALENT AREA

TOTAL

LIFT

VOLUME

Series8

Series11

DP/P COMPARISON

-0.050

-0.040

-0.030

-0.020

-0.010

0.000

0.010

0.020

0.030

0.040

0 20 40 60 80 100 120 140 160 180 200 220 240

STA ~ X ~ FT

DP/P

BoomScriptDPP

SEEBFFUN

COMP DATA 1

COMP DATA 2

Series5
48 

c)

Figure 2.1.12 Boom summary of 765-070A, M1.6, 270,000lbs, 51,000ft, ~2.4 psf boom. a) Δp/p at 10 

body lengths, b) equivalent area distribution, c) ground pressure signature and PLdB from 

Zephyrus. 

 

DATATABLE 5

ZEPHYRUS PERCEIVED LOUDNESS RESULTS

OASPL DBA DBC PLDB

124.48 93.71 109.44 109.16

*EOD5

Total @ Receiver With Reflection

-2.8

-2.4

-2.0

-1.6

-1.2

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

2.0

2.4

2.8

-0.05 0.00 0.05 0.10 0.15 0.20 0.25 0.30

Time, S

Signature, Psf

RECEIVER_REFL
49 

Table 2.1.13 Characteristics data of 765-071B 

Figure 2.1.13 General arrangement of 765-071B. 

115.15ft 

214.38 ft

model DATA TABLE (as drawn airplane)

765-071B Wing Horizontal Canard Vertical

ITEM ESDU Total Tail Tail

Area to CL 4284.177 4699.1 254.40 136.98 664.42

Exposed 254.40 136.98 559.18

Reference 4284.177 254.40 136.98 559.18

Aspect Ratio 3.09 3.43 3.15 3.15 1.431

Taper Ratio 0.084754 - 0.25 3.150 0.229

LE Sweep angle 0.00 48 48 45.54

Dihedral, TE 12 62.928 0 0 14

T/C 0.024 0.024 0.000 0.000

Tail Volume 0.136 0.068 0.068 0.1316

Span , in 1381.753 903.808 339.70 249.27 240.01

Root Chord, In 823.1868 788.66 172.55 132.14 273.87

Tip Chord, in 69.768 56.000 43.14 26.13 62.73

M.A.C. IN 552.4253 761.020 120.78 90.97 190.38

X 1/4 mac 1809.845 1653.692 2446.45 640.50 2366.73

Y, Zmac 248.2853 330.44368 104.370 126.91

Tail Arm, IN. 636.61 1169.34 556.8896
50 

Figure 2.1.14 Features of 765-071B. 

 

•Design TOGW= 300,000 lbs 

•Design Cruise Mach=1.8 

•Sref =4284 sqft 

•Design Range (100-106 pass.)= 4000 nmi 

•Cruise L/D=8.73 * 

•60K lb thrust class “BSE2” engine 

•Fan diam. =84” 

•HSCT-like LE, TE devices 

•3-surface pitch control /trim/ ride qual. 

•Fuel burn 0.28 lb/seat nmi @ 100 pass. 

 (= 3.6 seat nmi /lb) 

*L/D=8.02 on -071A with lower-risk aft t/c 

 but higher sized TOGW and boom level 
51 

Table 2.1.15 Weights summary of 765-071B. 
52 

Figure 2.1.16 Interior arrangement of 765-071B. 
53 

DP/P COMPARISON

-0.060

-0.050

-0.040

-0.030

-0.020

-0.010

0.000

0.010

0.020

0.030

0.040

0 20 40 60 80 100 120 140 160 180 200 220 240

STA ~ X ~ FT

DP/P

BoomScriptDPP

SEEBFFUN

COMP DATA 1

COMP DATA 2

Series5

EQUIVALENT AREA COMPARISON

-10

90

190

290

390

490

0 20 40 60 80 100 120 140 160 180 200 220 240

STA ~ X ~ FT

EQUIVALENT AREA

TOTAL

LIFT

VOLUME

Series8

Series11

a) 

b) 
54 

c) 

Figure 2.1.17 Boom assessment of 765-071B at initial cruise M=1.6, 270,000lbs, 51,000 ft, ~1.7 psf 

boom. a) Δp/p at 10 body lengths, b) equivalent area distribution, c) ground signature and PLdB 

from Zephyrus. 

Ground Signature

-2.0

-1.8

-1.6

-1.4

-1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

0.10 0.15 0.20 0.25 0.30 0.35 0.40 0.45

Time, S

Pressure, psf

PLOTPHI
55 

Figure 2.2.1.1 Pareto optima with the Mach number for T080 area-ruling (M_design) either fixed 

(red) or variable (brown). Optimization maximized cruise L/D and total body volume as objectives. 

Mdesign free

Mdesign = Mcruise
56 

Figure 2.2.1.2. Pareto optima maximizing L/D and body volume. L/D alternatively at cruise (green), 

at the thrust pinch Mach 1.1 (blue), and a weighted sum (red) of cruise and pinch. It is important to 

consider the L/D at thrust pinch, and doing so will affect the L/D at cruise. 

Figure 2.2.1.3. Pareto optima maximizing L/D and body volume for various depth of aft body at the 

wing box. Adding depth reduced L/D. 

increasing 

fuselage depth 

at spar box 

cost of optimizing for cruise alone 

cost of not optimizing for cruise alone 
57 

Figure 2.2.1.4 Body area distributions from increasing fuselage depth at the rear spar, for maximum 

L/D at fixed volume. 

Figure 2.2.1.5 Effect of outboard wing span and sweep on boom PLdB. Only the outboard panel 

varied, but the span fraction represents the ratio of overall span to original overall span. 
58 

Figure 2.2.1.6 Effect of varying cruise altitude (ft) on boom PLdB for various Mach. Mach 1.8 is 

maximum. Jump near 55,000ft corresponds to coalescence of double shock. 

Figure 2.2.1.7 Effect of varying altitude on boom signature for Mach 1.8. Favorable double shock at 

the front coalesces and is lost for higher cruise altitude. 
59 

Figure 2.2.1.8 Pareto optima: (+) 2 objectives, maximize L/D & maximize volume; (x) 3 objectives, 

maximize L/D, maximize volume, minimize PLdB. The frontier of L/D & volume for each is 

essentially the same. 
60 

Figure 2.2.1.9 Pareto optima: (+) 2 objectives, maximize L/D & maximize volume with no influence 

of PLdB; (x) 3 objectives, includes minimizing PLdB. Lines are contours of constant PLdB (105.2 to 

107.4 in steps of 0.2) of a best fit approximation to the data. 2-objective optima vary in PLdB, and 3-

objective optima show that the cause is a steep trade in PLdB near the 2-objective L/D-volume 

frontier. 

Figure 2.2.1.10 Pareto optima, maximizing volume & L/D while minimizing PLdB. Curves are cuts through best-fit 

surface of the Pareto frontier. 
61 

a) 

b) c) 

Figure 2.2.1.11a, b & c. Orthogonal views of 3-objective Pareto; maximize L/D, maximize bodyvol, 

minimize PLdB. Aft PLdB is the dominant contributor to overall PLdB. Additional degrees of 

freedom in forward and aft body result in a similar trade between L/D and volume at their frontier. 
62 

Figure 2.2.1.12 Initial N+2 Concept: 765-072B 
63 

Figure 2.2.2.1A. Supersonic Non-Stop Mission Profile 

Boeing Mission Rules

z Nominal Performancez Standard Dayz Fuel Density: 6.7 lb/US Gallon

Mission

Still Air Range

Block Time & Fuel

Climb

and

Accelerate

to

Optimum

Supersonic

Cruise

M

&

h

Climbout and Accelerate

to 1,500 ft & 250 kts

Taxi Out (9 Minutes)

Takeoff to 35 ft

Approach and Land (4 Minutes)

Taxi In (5 Minutes from Reserves)

SupersonicCruise / ClimbConstant Mach, Optimum Altitude

(55k ft limit)

See specified 

climb / descent 

Schedule to right

See climb schedules on next chart

Descend

and

decelerate

Reserves

Zero Wind

200 nm

Economy Climb

Cruise at LRC Mach

Approach & Land (4 Minutes)

5% Block Fuel Allowance

Missed Approach

Descend to 1,500 ft

Boeing Mission Rules

z Nominal Performancez Standard Dayz Fuel Density: 6.7 lb/US Gallon

Mission

Still Air Range

Block Time & Fuel

Climb

and

Accelerate

to

Optimum

Supersonic

Cruise

M

&

h

Climbout and Accelerate

to 1,500 ft & 250 kts

Taxi Out (9 Minutes)

Takeoff to 35 ft

Approach and Land (4 Minutes)

Taxi In (5 Minutes from Reserves)

SupersonicCruise / ClimbConstant Mach, Optimum Altitude

(55k ft limit)

See specified 

climb / descent 

Schedule to right

See climb schedules on next chart

Descend

and

decelerate

Reserves

Zero Wind

200 nm

Economy Climb

Cruise at LRC Mach

Approach & Land (4 Minutes)

5% Block Fuel Allowance

Missed Approach

Descend to 1,500 ft

200 nm

Economy Climb

Cruise at LRC Mach

Approach & Land (4 Minutes)

5% Block Fuel Allowance

Missed Approach

Descend to 1,500 ft
64 

Figure 2.2.2.1B. Supersonic Non-Stop Mission Profile (con’t) 

Boeing Mission Rules

z Nominal Performancez Standard Dayz Fuel Density: 6.7 lb/US Gallon

Climb / Descent Schedule

Mission ALTITUDE (FT) MACH 

Profile Initial Final Initial Final Notes 

CLIMB 1500 10000 0.39 0.45 Climb at 250 KCAS from 1500 ft to 10,000 ft CLIMB 10000 10300 0.45 0.69 Accelerate to 375 KEAS at roughly constant 

altitude (10,300 ft) 

CLIMB 10300 20774 0.69 0.85 Climb at constant 375 KEAS to Mach 0.85 

(~20,800 ft altitude) 

CLIMB 20774 35000 0.85 0.85 Climb at constant Mach 0.85 to 35,000 ft CLIMB 35000 39000 0.85 0.95 Climb and accelerate to Mach 0.95 at 39,000 ft CLIMB 39000 41000 0.95 M-crz Climb and accelerate to Supersonic Cruise 

Mach (1.6 to 2.0) at 41,000 ft 

CLIMB 41000 h-opt M-crz M-crz Climb to optimum initial cruise altitude 

 Climb / cruise with 53,000 ft maximum altitude 

DESCENT 53000 39000 M-crz 0.95 Descend & decelerate to Mach 0.95 @ 39,000 

ft 

DESCENT 39000 34960 0.95 0.85 Descend to Mach 0.85 and 273 KEAS (altitude 

~ 35,000 ft) 

DESCENT 34960 20774 0.85 0.85 Descend at Mach 0.85 and 375 KEAS (altitude 

~20,800 ft) 

DESCENT 20774 10300 0.85 0.69 Descend at constant 375 KEAS to ~10,300 ft DESCENT 10300 10000 0.69 0.45 Decelerate to 250 KCAS at roughly constant 

altitude (10,000 ft) 

DESCENT 10000 1500 0.45 0.39 Descend to 1500 ft at constant 250 KCAS 

Climb/Cruise with 55,000 ft maximum altitude

Boeing Mission Rules

z Nominal Performancez Standard Dayz Fuel Density: 6.7 lb/US Gallon

Climb / Descent Schedule

Mission ALTITUDE (FT) MACH 

Profile Initial Final Initial Final Notes 

CLIMB 1500 10000 0.39 0.45 Climb at 250 KCAS from 1500 ft to 10,000 ft CLIMB 10000 10300 0.45 0.69 Accelerate to 375 KEAS at roughly constant 

altitude (10,300 ft) 

CLIMB 10300 20774 0.69 0.85 Climb at constant 375 KEAS to Mach 0.85 

(~20,800 ft altitude) 

CLIMB 20774 35000 0.85 0.85 Climb at constant Mach 0.85 to 35,000 ft CLIMB 35000 39000 0.85 0.95 Climb and accelerate to Mach 0.95 at 39,000 ft CLIMB 39000 41000 0.95 M-crz Climb and accelerate to Supersonic Cruise 

Mach (1.6 to 2.0) at 41,000 ft 

CLIMB 41000 h-opt M-crz M-crz Climb to optimum initial cruise altitude 

 Climb / cruise with 53,000 ft maximum altitude 

DESCENT 53000 39000 M-crz 0.95 Descend & decelerate to Mach 0.95 @ 39,000 

ft 

DESCENT 39000 34960 0.95 0.85 Descend to Mach 0.85 and 273 KEAS (altitude 

~ 35,000 ft) 

DESCENT 34960 20774 0.85 0.85 Descend at Mach 0.85 and 375 KEAS (altitude 

~20,800 ft) 

DESCENT 20774 10300 0.85 0.69 Descend at constant 375 KEAS to ~10,300 ft DESCENT 10300 10000 0.69 0.45 Decelerate to 250 KCAS at roughly constant 

altitude (10,000 ft) 

DESCENT 10000 1500 0.45 0.39 Descend to 1500 ft at constant 250 KCAS 

Climb/Cruise with 55,000 ft maximum altitude

Mission ALTITUDE (FT) MACH 

Profile Initial Final Initial Final Notes 

CLIMB 1500 10000 0.39 0.45 Climb at 250 KCAS from 1500 ft to 10,000 ft CLIMB 10000 10300 0.45 0.69 Accelerate to 375 KEAS at roughly constant 

altitude (10,300 ft) 

CLIMB 10300 20774 0.69 0.85 Climb at constant 375 KEAS to Mach 0.85 

(~20,800 ft altitude) 

CLIMB 20774 35000 0.85 0.85 Climb at constant Mach 0.85 to 35,000 ft CLIMB 35000 39000 0.85 0.95 Climb and accelerate to Mach 0.95 at 39,000 ft CLIMB 39000 41000 0.95 M-crz Climb and accelerate to Supersonic Cruise 

Mach (1.6 to 2.0) at 41,000 ft 

CLIMB 41000 h-opt M-crz M-crz Climb to optimum initial cruise altitude 

 Climb / cruise with 53,000 ft maximum altitude 

DESCENT 53000 39000 M-crz 0.95 Descend & decelerate to Mach 0.95 @ 39,000 

ft 

DESCENT 39000 34960 0.95 0.85 Descend to Mach 0.85 and 273 KEAS (altitude 

~ 35,000 ft) 

DESCENT 34960 20774 0.85 0.85 Descend at Mach 0.85 and 375 KEAS (altitude 

~20,800 ft) 

DESCENT 20774 10300 0.85 0.69 Descend at constant 375 KEAS to ~10,300 ft DESCENT 10300 10000 0.69 0.45 Decelerate to 250 KCAS at roughly constant 

altitude (10,000 ft) 

DESCENT 10000 1500 0.45 0.39 Descend to 1500 ft at constant 250 KCAS 

Climb/Cruise with 55,000 ft maximum altitude
65 

Figure 2.2.3.1 Aerodynamic Differences between -070 and -072B Species 

• Primary configuration differences

– Horizontal tail on -070– Different planform (sweeps, thickness, aspect ratio)

L/D @ M 1.6 Comparison

Sgross=5350, Fn=84000 lbs, STF-1336-01

02468

1012

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

CL

L/D

072B Type070 Type

• Primary configuration differences

– Horizontal tail on -070– Different planform (sweeps, thickness, aspect ratio)

L/D @ M 1.6 Comparison

Sgross=5350, Fn=84000 lbs, STF-1336-01

02468

1012

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

CL

L/D

072B Type070 Type
66 

Figure 2.2.3.2 Partial Power Takeoff for 1100 fps Vj 

Case 1: Engine sized for performance also meets 10,000 ft takeoff at 1100 fps Vj

0

2000400060008000

1000012000140001600018000

50% 60% 70% 80% 90% 100%

% Max Takeoff Thrust

Balanced Field Distance (ft)

500700900110013001500

Jet Velocity (fps) Dista

Vj

-072B Type SpeciesCompromise #1 EngineMach 1.6, 4000 nm Range

10% FAR

allowance

Sized for 

Case 1: Engine sized for performance also meets 10,000 ft takeoff at 1100 fps Vj

0

2000400060008000

1000012000140001600018000

50% 60% 70% 80% 90% 100%

% Max Takeoff Thrust

Balanced Field Distance (ft)

500700900110013001500

Jet Velocity (fps) Dista

Vj

-072B Type SpeciesCompromise #1 EngineMach 1.6, 4000 nm Range

10% FAR

allowance

Sized for 

0

2000400060008000

1000012000140001600018000

50% 60% 70% 80% 90% 100%

% Max Takeoff Thrust

Balanced Field Distance (ft)

500700900110013001500

Jet Velocity (fps) Dista

Vj

-072B Type SpeciesCompromise #1 EngineMach 1.6, 4000 nm Range

10% FAR

allowance

Sized for 
67 

Figure 2.2.3.3 Sizing for 1100 fps Vj at Takeoff 

Case 2: Engine must be sized up to meet 10,000 ft takeoff at 1100 fps Vj

0

2000400060008000

100001200014000160001800020000

50% 60% 70% 80% 90% 100%

% Max Takeoff Thrust

Balanced Field Distance (ft)

5007009001100130015001700

Jet Velocity (fps) Dista

DistaVj

-072B Type SpeciesLow VJ #2 EngineMach 1.6, 4000 nm Range

10% FAR

allowance

Sized f

Sized fo

Thrust

Increase

1180 fps Vj

With FAR allowance

Case 2: Engine must be sized up to meet 10,000 ft takeoff at 1100 fps Vj

0

2000400060008000

100001200014000160001800020000

50% 60% 70% 80% 90% 100%

% Max Takeoff Thrust

Balanced Field Distance (ft)

5007009001100130015001700

Jet Velocity (fps) Dista

DistaVj

-072B Type SpeciesLow VJ #2 EngineMach 1.6, 4000 nm Range

10% FAR

allowance

Sized f

Sized fo

Thrust

Increase

1180 fps Vj

With FAR allowance

0

2000400060008000

100001200014000160001800020000

50% 60% 70% 80% 90% 100%

% Max Takeoff Thrust

Balanced Field Distance (ft)

5007009001100130015001700

Jet Velocity (fps) Dista

DistaVj

-072B Type SpeciesLow VJ #2 EngineMach 1.6, 4000 nm Range

10% FAR

allowance

Sized f

Sized fo

Thrust

Increase

1180 fps Vj

With FAR allowance
68 

Figure 2.2.3.4 072B Type Species Carpet Plot 

FOM

2.603.304.004.70

Fn

_Base=45000 Fn

_Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_ft=3000

sgross_ft=4000

sgross_ft=

300 fpmROC 10K ft Bal Field Dist @60%pwr WingFuel Only 78psf W/S Selected Point Wing area= 4600 ft2

Thrust = 55,000 lbs

FOM = 3.7

Mach 1.6, 4000 nm, 100 PAX,Compromise #1 Engine

Mach 1.6, 4000 nm, 100 PAX,Compromise #1 Engine

FOM

2.603.304.004.70

Fn

_Base=45000 Fn

_Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_ft=3000

sgross_ft=4000

sgross_ft=

300 fpmROC 10K ft Bal Field Dist @60%pwr WingFuel Only 78psf W/S Selected Point Wing area= 4600 ft2

Thrust = 55,000 lbs

FOM = 3.7

Mach 1.6, 4000 nm, 100 PAX,Compromise #1 Engine

Mach 1.6, 4000 nm, 100 PAX,Compromise #1 Engine
69 

Figure 2.2.3.5 070 Type Species Carpet Plot 

FOM

2.603.304.004.70

Fn

_Base=45000

Fn_Base=50000

Fn_Base=60000

Fn_Base=70000

Base=80000

sgross=3000

sgross=4000

sgross=5000

300

fpm

ROC

10K ft Bal Field Dist @60%pwr WingFuel On W ly ing&Fuse

Fuel

78

psf

W/S

Selected Point

Wing area= 4550 ft2

Thrust = 59,500 lbs

FOM = 3.36

Mach 1.6, 4000 nm, 100 PAX,Compromise #1 Engine

Mach 1.6, 4000 nm, 100 PAX,Compromise #1 Engine

FOM

2.603.304.004.70

Fn

_Base=45000

Fn_Base=50000

Fn_Base=60000

Fn_Base=70000

Base=80000

sgross=3000

sgross=4000

sgross=5000

300

fpm

ROC

10K ft Bal Field Dist @60%pwr WingFuel On W ly ing&Fuse

Fuel

78

psf

W/S

Selected Point

Wing area= 4550 ft2

Thrust = 59,500 lbs

FOM = 3.36

Mach 1.6, 4000 nm, 100 PAX,Compromise #1 Engine

Mach 1.6, 4000 nm, 100 PAX,Compromise #1 Engine
70 

Figure 2.2.3.6 Species Trade Results 

0

10000200003000040000500006000070000

070 Type 

Installed Thrust per Engine (lbs)

0

50000

100000150000200000250000300000350000

070 Type 

TOGW (lbs)

012345

070 Type 072B Type

FOM

100 PAXCompromise #1 EngineMach 1.6, 4000 nm Range

••Select 072B species for rest of trades Select 072B species for rest of trades

Goal

0

10000200003000040000500006000070000

070 Type 

Installed Thrust per Engine (lbs)

0

50000

100000150000200000250000300000350000

070 Type 

TOGW (lbs)

012345

070 Type 072B Type

FOM

100 PAXCompromise #1 EngineMach 1.6, 4000 nm Range

••Select 072B species for rest of trades Select 072B species for rest of trades

Goal
71 

Figure 2.2.4.1. Passenger Layout for Number of Passengers Trade 

• Passengers laid out per IAC Short / 

Medium Dual Class rules, as BCA stud

• Three “nominal” cross-sections develo

BLUE RED

GOL

• Passengers laid out per IAC Short / 

Medium Dual Class rules, as BCA stud

• Three “nominal” cross-sections develo

BLUE RED

GOL
72 

Figure 2.2.4.2 Payload and Fuselage Sizing 

• 100 Pax – 2+2 and 2+1, 2 x LD-4 baggage• 50 Pax – 2+1 and 1+1, 1 x LD-4 baggage• 25 Pax – 2+1 and 1+1, Bulk baggage

2+1 1+1

OAL 185

2+1 1+1

OAL 190

2+2 2+1

OAL 

Fineness Ratio (l/d) ~ 2

• 100 Pax – 2+2 and 2+1, 2 x LD-4 baggage• 50 Pax – 2+1 and 1+1, 1 x LD-4 baggage• 25 Pax – 2+1 and 1+1, Bulk baggage

2+1 1+1

OAL 185

2+1 1+1

OAL 190

2+2 2+1

OAL 

Fineness Ratio (l/d) ~ 2
73 

Figure 2.2.4.3 Carpet Plot for 50 Passenger Concept 

FOM

1.501.802.102.402.70

Fn

_Base=40000

Fn

_Base=50000

Fn_Base=60000

sgross_ft=3000

sgross

_ft=4000

sgross

_ft=5000

300 fpm ROC 10K Bal Field Dist @ 60% pwr Wing

Fuel Only

Wing&Fuse

Fuel

78 psf W/S

Selected Point

Wing area= 3500 ft2

Thrust = 43,000 lbs

FOM = 2.3

Mach 1.6, 4000 nm, Compromise #1 Engine Mach 1.6, 4000 nm, Compromise #1 Engine

(-24%)(-22%)

FOM

1.501.802.102.402.70

Fn

_Base=40000

Fn

_Base=50000

Fn_Base=60000

sgross_ft=3000

sgross

_ft=4000

sgross

_ft=5000

300 fpm ROC 10K Bal Field Dist @ 60% pwr Wing

Fuel Only

Wing&Fuse

Fuel

78 psf W/S

Selected Point

Wing area= 3500 ft2

Thrust = 43,000 lbs

FOM = 2.3

FOM

1.501.802.102.402.70

Fn

_Base=40000

Fn

_Base=50000

Fn_Base=60000

sgross_ft=3000

sgross

_ft=4000

sgross

_ft=5000

300 fpm ROC 10K Bal Field Dist @ 60% pwr Wing

Fuel Only

Wing&Fuse

Fuel

78 psf W/S

Selected Point

Wing area= 3500 ft2

Thrust = 43,000 lbs

FOM = 2.3

Mach 1.6, 4000 nm, Compromise #1 Engine Mach 1.6, 4000 nm, Compromise #1 Engine

(-24%)(-22%)
74 

Figure 2.2.4.4 Carpet Plot for 100 Passenger Concept 

FOM

0.851.051.251.451.65

Fn

_Base=35000

Fn

_Base=45000

Fn

_

Base=55000

sgross_

ft=2500

sgross

_ft=3500

sgross

_ft=4

300 fpm ROC 10K ft Bal Field Dist 

Wing Fuel Only Wing& Fuse Fuel 78 psf W/S

Mach 1.6, 4000 nm, Compromise #1 Engine Mach 1.6, 4000 nm, Compromise #1 Engine

Selected Point

Wing area= 3050 ft2

Thrust = 37,500 lbs

FOM = 1.3

FOM

0.851.051.251.451.65

Fn

_Base=35000

Fn

_Base=45000

Fn

_

Base=55000

sgross_

ft=2500

sgross

_ft=3500

sgross

_ft=4

300 fpm ROC 10K ft Bal Field Dist 

Wing Fuel Only Wing& Fuse Fuel 78 psf W/S

Mach 1.6, 4000 nm, Compromise #1 Engine Mach 1.6, 4000 nm, Compromise #1 Engine

Selected Point

Wing area= 3050 ft2

Thrust = 37,500 lbs

FOM = 1.3
75 

Figure 2.2.4.5 Passenger Trade Results Summary 

0

0.5

1

1.5

2

2.5

3

3.5

4

25 50 100

Number of Passengers

FOM0

50000

100000150000200000250000300000

25 50 

Number of Pass

TOGW

0

100002000030000400005000060000

25 50 

Number of Passe

Installed Thrust per engine (lbs)

Mach 1.6, 4000 nm, Compromise #1 Engine Mach 1.6, 4000 nm, Compromise #1 Engine

•• 100 PAX is only option that exceeds FOM goal 100 PAX is only option that exceeds FOM goal

Goal

(-21%) (-32%)

0

0.5

1

1.5

2

2.5

3

3.5

4

25 50 100

Number of Passengers

FOM0

50000

100000150000200000250000300000

25 50 

Number of Pass

TOGW

0

100002000030000400005000060000

25 50 

Number of Passe

Installed Thrust per engine (lbs)

Mach 1.6, 4000 nm, Compromise #1 Engine Mach 1.6, 4000 nm, Compromise #1 Engine

•• 100 PAX is only option that exceeds FOM goal 100 PAX is only option that exceeds FOM goal

Goal

(-21%) (-32%)
76 

Figure 2.2.5.1 Carpet Plot for Mach 1.8 Concept 

FOM

2.603.304.004.70

Fn

_Base=50000

Fn

_

Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_

ft=3000

sgross_ft=4000

sgross

_ft=5000

300

fpm

ROC

10K ft Bal Field Dist @ 60% pwr Wing Fuel Only

Wing&Fuse

Fuel

78 psf W/S

4000 nm, 100 PAX, Compromise #1 Engine 4000 nm, 100 PAX, Compromise #1 Engine

Selected Point

Wing area= 4700 ft2

Thrust = 57,500 lbs

FOM = 3.55

(+2%)(+5%)

(-4%)

FOM

2.603.304.004.70

Fn

_Base=50000

Fn

_

Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_

ft=3000

sgross_ft=4000

sgross

_ft=5000

300

fpm

ROC

10K ft Bal Field Dist @ 60% pwr Wing Fuel Only

Wing&Fuse

Fuel

78 psf W/S

4000 nm, 100 PAX, Compromise #1 Engine 4000 nm, 100 PAX, Compromise #1 Engine

Selected Point

Wing area= 4700 ft2

Thrust = 57,500 lbs

FOM = 3.55

(+2%)(+5%)

(-4%)
77 

Figure 2.2.5.2 Carpet Plot for Mach 2.0 Concept 

FOM

1.602.102.603.103.60

Fn

_Base=65000 Fn_Base=75000 Fn

_Base=85000 sgross_ft=4000

sgross_ft=5000

sgross_ft=6000

300 fpm ROC

10K ft Bal Field Dist @ 60% p

Wing Only Fuel

78

psf

W/S

4000 nm, 100 PAX, Compromise #1 Engine 4000 nm, 100 PAX, Compromise #1 Engine

Selected Poin

Wing area= 

Thrust = 77

FOM = 2.55

(-30%)

FOM

1.602.102.603.103.60

Fn

_Base=65000 Fn_Base=75000 Fn

_Base=85000 sgross_ft=4000

sgross_ft=5000

sgross_ft=6000

300 fpm ROC

10K ft Bal Field Dist @ 60% p

Wing Only Fuel

78

psf

W/S

4000 nm, 100 PAX, Compromise #1 Engine 4000 nm, 100 PAX, Compromise #1 Engine

Selected Poin

Wing area= 

Thrust = 77

FOM = 2.55

(-30%)
78 

Figure 2.2.5.3 Cruise Mach Number Trade Results 

01234

1.6 1.8 2

Mach Number

FOM0 50000

100000150000200000250000300000350000400000

1.6 1.8 

Mach Num

TOGW

0

100002000030000400005000060000700008000090000

1.6 1.8 

Mach Num

Installed Thrust per Engine (lbs)

4000 nm, 100 PAX 4000 nm, 100 PAX

• Little difference between Mach 1.6 and 1.8• Other considerations (productivity) could favor M1.8• Other studies have shown M1.8 to have best FOM

• Little difference between Mach 1.6 and 1.8• Other considerations (productivity) could favor M1.8• Other studies have shown M1.8 to have best FOM

Goal

01234

1.6 1.8 2

Mach Number

FOM0 50000

100000150000200000250000300000350000400000

1.6 1.8 

Mach Num

TOGW

0

100002000030000400005000060000700008000090000

1.6 1.8 

Mach Num

Installed Thrust per Engine (lbs)

4000 nm, 100 PAX 4000 nm, 100 PAX

• Little difference between Mach 1.6 and 1.8• Other considerations (productivity) could favor M1.8• Other studies have shown M1.8 to have best FOM

• Little difference between Mach 1.6 and 1.8• Other considerations (productivity) could favor M1.8• Other studies have shown M1.8 to have best FOM

Goal
79 

Figure 2.2.6.1 Carpet Plot for 5000 nm Range Concept 

FOM

1.702.202.703.203.70

Fn

_

Base=70000

Fn_Base=80000

Fn_Base=90000

sgross_

ft=4000

sgross_ft=5000

sgross_ft=6000

sgross

_ft=

300 fpm ROC 10K ft Bal Field Dist @ 60%

pwr

Wing Fuel Only

Wing&Fuse Fuel

78 psf W/S

Mach 1.6, 100 PAX, Compromise #1 Engine Mach 1.6, 100 PAX, Compromise #1 Engine

Selected Point

Wing area= 6050 ft2

Thrust = 73,000 lbs

FOM = 2.95

(+32%)(+33%)(range +25%)

FOM

1.702.202.703.203.70

Fn

_

Base=70000

Fn_Base=80000

Fn_Base=90000

sgross_

ft=4000

sgross_ft=5000

sgross_ft=6000

sgross

_ft=

300 fpm ROC 10K ft Bal Field Dist @ 60%

pwr

Wing Fuel Only

Wing&Fuse Fuel

78 psf W/S

Mach 1.6, 100 PAX, Compromise #1 Engine Mach 1.6, 100 PAX, Compromise #1 Engine

Selected Point

Wing area= 6050 ft2

Thrust = 73,000 lbs

FOM = 2.95

(+32%)(+33%)(range +25%)
80 

Figure 2.2.6.2 Carpet Plot for 6000 nm Range Concept 

FOM

2.002.252.502.753.00

Fn

_Base=80000

Fn_Base=90000

Fn_Base=100000

sgross_

ft=5000

sgross_ft=6000

sgross

_ft=7

300 fpm ROC

Wing

Fuel

Only

Wing&Fuse

Fuel

All points do not meet 10,000 ft balanced field distance at 60% pow

All points do not meet 78 psf wing loading

Mach 1.6, 100 PAX, Compromise #1 Engine Mach 1.6, 100 PAX, Compromise #1 Engine

Selected Point

Wing area= 6100 ft2

Thrust = 90,500 lbs

FOM = 2.55

FOM

2.002.252.502.753.00

Fn

_Base=80000

Fn_Base=90000

Fn_Base=100000

sgross_

ft=5000

sgross_ft=6000

sgross

_ft=7

300 fpm ROC

Wing

Fuel

Only

Wing&Fuse

Fuel

All points do not meet 10,000 ft balanced field distance at 60% pow

All points do not meet 78 psf wing loading

Mach 1.6, 100 PAX, Compromise #1 Engine Mach 1.6, 100 PAX, Compromise #1 Engine

Selected Point

Wing area= 6100 ft2

Thrust = 90,500 lbs

FOM = 2.55
81 

Figure 2.2.6.3 Range Trade Results 

01234

4000 5000 6000

Range (nm)

FOM

*

* 6000 nm case does not meet W/S and

balanced field constraints

0

100000200000300000400000500000600000

4000 5000 6

Range (nm)

TOGW

0

20000400006000080000

100000

4000 5000 6

Range (nm)

Installed Thrust per Engine (lbs)

Mach 1.6, 100 PAX Mach 1.6, 100 PAX

Goal

01234

4000 5000 6000

Range (nm)

FOM

*

* 6000 nm case does not meet W/S and

balanced field constraints

0

100000200000300000400000500000600000

4000 5000 6

Range (nm)

TOGW

0

20000400006000080000

100000

4000 5000 6

Range (nm)

Installed Thrust per Engine (lbs)

Mach 1.6, 100 PAX Mach 1.6, 100 PAX

Goal
82 

Figure 2.2.7.1 Combined Range, Mach, Passenger Trade Results 

01234

3000 4000 5000 6000 7000

Range (nm)

FOM100 PAX

50 PAX25 PAX

072B Type SpeciesCompromise #1 EngineMach 2.0

01234

3000 4000 5000 6000 7000

Range (nm)

FOM

072B Type SpecieCompromise #1 EMach 1.8

01234

3000 4000 5000 6000 7000

Range (nm)

FOM100 PAX

50 PAX25 PAX

072B Type SpeciesCompromise #1 EngineMach 1.6

• 4000 nm and 100 PAX exceeds FOM goal at Mach 1.6 and 

• 4000 nm and 100 PAX exceeds FOM goal at Mach 1.6 and 

Does not meet W/S constrain

Goal

Goal

Goal

01234

3000 4000 5000 6000 7000

Range (nm)

FOM100 PAX

50 PAX25 PAX

072B Type SpeciesCompromise #1 EngineMach 2.0

01234

3000 4000 5000 6000 7000

Range (nm)

FOM

072B Type SpecieCompromise #1 EMach 1.8

01234

3000 4000 5000 6000 7000

Range (nm)

FOM100 PAX

50 PAX25 PAX

072B Type SpeciesCompromise #1 EngineMach 1.6

• 4000 nm and 100 PAX exceeds FOM goal at Mach 1.6 and 

• 4000 nm and 100 PAX exceeds FOM goal at Mach 1.6 and 

Does not meet W/S constrain

Goal

Goal

Goal
83 

Figure 2.2.8.1 Carpet Plot for Low Jet Velocity #1 Engine 

FOM

2.803.303.804.304.80

Fn

_

Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_

ft=3000

sgross_ft=4000

sgross

_ft=5000

300 fpm ROC

Wing

Fuel

Only

Wing&Fuse

Fu

el

7

8

p

s

f

W/

S

Selected P

Wing ar

Thrust = 

FOM = 4.0

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

All points meet 10,000 ft balanced field distance at 79% power

FOM

2.803.303.804.304.80

Fn

_

Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_

ft=3000

sgross_ft=4000

sgross

_ft=5000

300 fpm ROC

Wing

Fuel

Only

Wing&Fuse

Fu

el

7

8

p

s

f

W/

S

Selected P

Wing ar

Thrust = 

FOM = 4.0

FOM

2.803.303.804.304.80

Fn

_

Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_

ft=3000

sgross_ft=4000

sgross

_ft=5000

300 fpm ROC

Wing

Fuel

Only

Wing&Fuse

Fu

el

7

8

p

s

f

W/

S

Selected P

Wing ar

Thrust = 

FOM = 4.0

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

All points meet 10,000 ft balanced field distance at 79% power
84 

Figure 2.2.8.2 Carpet Plot for Low Jet Velocity #2 Engine 

FOM

2.202.703.203.704.20

Fn_Base=50000

Fn_Base=60000

Fn_

Base=70000

Fn_Base=80000

sgross_

ft=3000

sgross_

ft=4000

sgross_ft=5000

sgross_ft=5500

300 fpm

ROC

10K ft Bal Field Dist @ 66% pwr

Wing

Fuel Only

Wing&Fuse

Fuel

78 psf W/S

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

Selected Point

Wing area= 5400 ft2

Thrust = 70,000 lbs

FOM = 3.21

Sized forPerformance

FOM

2.202.703.203.704.20

Fn_Base=50000

Fn_Base=60000

Fn_

Base=70000

Fn_Base=80000

sgross_

ft=3000

sgross_

ft=4000

sgross_ft=5000

sgross_ft=5500

300 fpm

ROC

10K ft Bal Field Dist @ 66% pwr

Wing

Fuel Only

Wing&Fuse

Fuel

78 psf W/S

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

Selected Point

Wing area= 5400 ft2

Thrust = 70,000 lbs

FOM = 3.21

Sized forPerformance
85 

Figure 2.2.8.3 Carpet Plot for Compromise #1 Engine 

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

FOM

2.603.304.004.70

Fn

_Base=45000 Fn

_Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_ft=3000

sgross_ft=4000

sgross_ft=

300 fpmROC 10Kft Bal Field Dist @ 60%pw W r ingFuel Only 78psf W/S Selected Point Wing area= 4500 ft2

Thrust = 53,000 lbs

FOM = 3.7

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

FOM

2.603.304.004.70

Fn

_Base=45000 Fn

_Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_ft=3000

sgross_ft=4000

sgross_ft=

300 fpmROC 10Kft Bal Field Dist @ 60%pw W r ingFuel Only 78psf W/S Selected Point Wing area= 4500 ft2

Thrust = 53,000 lbs

FOMFOM = 3.7

2.603.304.004.70

Fn

_Base=45000 Fn

_Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_ft=3000

sgross_ft=4000

sgross_ft=

300 fpmROC 10Kft Bal Field Dist @ 60%pw W r ingFuel Only 78psf W/S Selected Point Wing area= 4500 ft2

Thrust = 53,000 lbs

FOM = 3.7
86 

Figure 2.2.8.4 Carpet Plot for Compromise #2 Engine 

FOM

2.102.603.103.604.104.60

Fn_Base=42000

Fn_Base=46000

Fn_Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_ft=3000

sgross_ft=4000

sgross

_ft=5000

300 fpm

ROC

Wing

Fuel

Only

Wing&Fuse

Fu

el

7

8

p

sf

W/S

All points do not meet 10,000 ft balanced field distance at 54% p

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

Selected Point

Wing area= 4450 ft2

Thrust = 48,500 lbs

FOM = 3.8

FOM

2.102.603.103.604.104.60

Fn_Base=42000

Fn_Base=46000

Fn_Base=50000

Fn_Base=60000

Fn_Base=70000

Fn_Base=80000

sgross_ft=3000

sgross_ft=4000

sgross

_ft=5000

300 fpm

ROC

Wing

Fuel

Only

Wing&Fuse

Fu

el

7

8

p

sf

W/S

All points do not meet 10,000 ft balanced field distance at 54% p

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

Selected Point

Wing area= 4450 ft2

Thrust = 48,500 lbs

FOM = 3.8
87 

Figure 2.2.8.5. Carpet Plot for Low Diameter #1 Engine 

FOM

2.202.703.203.704.204.70

Fn_Base=45000

Fn

_

Base=77500

Fn_Base=110000

sgross_

ft=4000

sgross_ft=4500

sgross

_ft=5000

300 fpm ROC

Wing Fuel Only 78 psf W/S

All points do not meet 10,000 ft balanced field distance at derate 

Selected Point

Wing area= 4300 ft2

Thrust = 49,500 lbs

FOM = 3.9

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

FOM

2.202.703.203.704.204.70

Fn_Base=45000

Fn

_

Base=77500

Fn_Base=110000

sgross_

ft=4000

sgross_ft=4500

sgross

_ft=5000

300 fpm ROC

Wing Fuel Only 78 psf W/S

All points do not meet 10,000 ft balanced field distance at derate 

Selected Point

Wing area= 4300 ft2

Thrust = 49,500 lbs

FOM = 3.9

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX
88 

Figure 2.2.8.6. Carpet Plot for Low Diameter #2 Engine 

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

FOM

2.002.503.003.504.004.50

Fn

_

Base=40000

Fn_Base=60000

Fn_Base=80000

sgross_

ft=3000

sgross_

ft=4000

sgross

_ft=5

300 fpm ROC

W

i

n

g

F

u

e

l

O

n

l

y

W

i

n

g

&

F

u

s

e

F

u

e

l

78

psf W/S

All points do not meet 10,000 ft balanced field distance at derate 

Selected Point

Wing area= 4550 

Thrust = 46,000 

FOM = 3.5

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

FOM

2.002.503.003.504.004.50

Fn

_

Base=40000

Fn_Base=60000

Fn_Base=80000

sgross_

ft=3000

sgross_

ft=4000

sgross

_ft=5

300 fpm ROC

W

i

n

g

F

u

e

l

O

n

l

y

W

i

n

g

&

F

u

s

e

F

u

e

l

78

psf W/S

All points do not meet 10,000 ft balanced field distance at derate 

Selected Point

Wing area= 4550 

Thrust = 46,000 

FOM = 3.5
89 

Figure 2.2.8.7 Engine Cycle Trade Results 

0

50000

100000150000200000250000300000350000

Low Vj #1Low Vj #2 Compromise #1 Compromise #2 Low Fan Dia #1 Low F

TOGW (lbs) 

0

1000020000300004000050000600007000080000

Low Vj #1Low Vj #2 Compromise #1 Compromise #2 Low Fan Dia #1 Low Fa

Installed Thrust per Engine (lbs) 

• Engine cycles are representative of different

technology levels and jet velocities

We are selecting Compromise #1 engine cycle for Task 2.5 We are selecting Compromise #1 engine cycle for Task 2.5

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

0

0.5

1

1.5

2

2.5

3

3.5

4

4.5

Low Vj #1Low Vj #2

Compromise #1 Compromise #2 Low Fan Dia #1 Low Fan Dia #2

FOMPerformance 1100 fps Vj

SIZING CRITERIA

1100 fps

1180 fps

1111 fps

1380 fps

1335 fps

1665 fps

Goal

0

50000

100000150000200000250000300000350000

Low Vj #1Low Vj #2 Compromise #1 Compromise #2 Low Fan Dia #1 Low F

TOGW (lbs) 

0

1000020000300004000050000600007000080000

Low Vj #1Low Vj #2 Compromise #1 Compromise #2 Low Fan Dia #1 Low Fa

Installed Thrust per Engine (lbs) 

• Engine cycles are representative of different

technology levels and jet velocities

We are selecting Compromise #1 engine cycle for Task 2.5 We are selecting Compromise #1 engine cycle for Task 2.5

Mach 1.6, 4000 nm, 100 PAX Mach 1.6, 4000 nm, 100 PAX

0

0.5

1

1.5

2

2.5

3

3.5

4

4.5

Low Vj #1Low Vj #2

Compromise #1 Compromise #2 Low Fan Dia #1 Low Fan Dia #2

FOMPerformance 1100 fps Vj

SIZING CRITERIA

1100 fps

1180 fps

1111 fps

1380 fps

1335 fps

1665 fps

Goal
90 

Figure 2.2.9.1 765-072B Performance Summary 

Configuration V2.1

Dec 17

Compromise 2

Wing Area (ESDU) 3527Wing Area (Gross) 4450Thrust per engine 48,500Engine T/W 5TOGW 273,834OEW 129,940Payload 21,000Total Fuel 123,217Block Fuel 105,725Range 4000Supersonic Cruise

Mach 1.6Altitude 52,245Weight 251,948CL 0.1875L/D 9.698SFC 1.0111

Climb (pinch point)

Mach 1.13Altitude 40,550Weight 261,221CL 0.2224L/D 10.142Thrust 27,125SFC 0.8533ROC 360

Bal Field Dist. (100% pwr) 6441W/S (takeoff) 77.6FOM 3.78

Note 1: V2.5 of the MDAO model has corrections to aero and engine sNote 2: Engines were not de-rated to meet 1100 fps takeoff VjNote 3: V2.5 cases were constrained to 55,000 ft max cruise altitude

V2.5 (resized)

Mar 12

Compromise 2

37714600

58,000

5

291,098135,320

21,000

135,130116,346

4000

1.6

52,412

267,223

0.1833

9.486

1.0107

1.13

40,550

277,651

0.2162

9.797

29,6410.8533

322

7088

77.23.44

Configuration V2.1

Dec 17

Compromise 2

Wing Area (ESDU) 3527Wing Area (Gross) 4450Thrust per engine 48,500Engine T/W 5TOGW 273,834OEW 129,940Payload 21,000Total Fuel 123,217Block Fuel 105,725Range 4000Supersonic Cruise

Mach 1.6Altitude 52,245Weight 251,948CL 0.1875L/D 9.698SFC 1.0111

Climb (pinch point)

Mach 1.13Altitude 40,550Weight 261,221CL 0.2224L/D 10.142Thrust 27,125SFC 0.8533ROC 360

Bal Field Dist. (100% pwr) 6441W/S (takeoff) 77.6FOM 3.78

Note 1: V2.5 of the MDAO model has corrections to aero and engine sNote 2: Engines were not de-rated to meet 1100 fps takeoff VjNote 3: V2.5 cases were constrained to 55,000 ft max cruise altitude

V2.5 (resized)

Mar 12

Compromise 2

37714600

58,000

5

291,098135,320

21,000

135,130116,346

4000

1.6

52,412

267,223

0.1833

9.486

1.0107

1.13

40,550

277,651

0.2162

9.797

29,6410.8533

322

7088

77.23.44
91 

Figure 2.2.9.2 765-072B External Views
92 

Figure 2.2.10.1 Features of the 765-078 boom softening airplanes. 

Cruciform tail 

for modulation 

of aft lift 

signature 

Wing Planform: 

IB LE sweep 78/68 deg 

OB LE sweep 44deg 

OB t/c 2.4% 

tip chord 95.8” 

Maintain Body Geometry Constraints : 

Length ~ 224’ 

Cockpit 60” wide at height of 54” above floor. 

Fwd cabin cross-sectional area 61 ft**2 (galley/lav/2-abreast econ). 

Aft-body cross-sectional area (APU/Vertical structural tie-in) 61 ft**2 at 

wing spar. 

>14,200 ft**3 total volume inside body OML. 

100 pax dual class LOPA ( <= 25 1st class pax). 

Goal: 

Variant of -072B with greater 

potential for “boom softening”

while maintaining efficiency 

Higher aspect 

ratio, higher 

sweep canard 

(~horizontal tail) 
93 

Figure 2.2.10.2 765-078A 3-view and characteristics. 

109.82ft 

229.57 ft

model DATA TABLE (as drawn airplane)

Model 765-078A Wing Horizontal Canard Vertical

ITEM ESDU Total Tail Tail

Area to CL 4016.675 4366.2 161.99 300.84 409.14

Exposed 161.99 300.84 387.78

Reference 4016.675 161.99 300.84 387.78

Aspect Ratio 3.00 3.43 3.15 3.15 2.000

Taper Ratio 0.122449 - 0.25 3.150 0.241

LE Sweep angle 0.00 48 48 58

Dihedral, TE 12 55 10 0 0

T/C 0.024 0.024 0.000 0.030

Tail Volume 0.205 0.090 0.115 0.1355

Span , in 1317.84 903.808 271.07 369.41 236.31

Root Chord, In 782.0418 788.66 137.69 187.63 391.70

Tip Chord, in 95.76 56.000 34.42 46.91 94.52

M.A.C. IN 528.3255 728.109 96.38 131.34 273.38

X 1/4 mac 1713.465 1563.924 540.31 2525.89 2455.02

Y, Zmac 243.6007 96.213896 79.881 136.08

Tail Arm, IN. 1173.16 812.42 741.5502
94 

Figure 2.2.10.3 MDboom and Zephyrus predictions of boom for the 765-078A, M=1.8, 270,000lb, 51,000
95 

Figure 2.2.10.4 MDboom and Zephyrus predictions of boom for the 765-078B, M=1.8, 270,000lb, 51,000
96 

Figure 2.3.1 765-076E 
97 

Figure 2.3.1.1 Early 765-076 study minimizing PLdB via body cross-section only (red +); and via body crosssection, body length, wing span and sweep (green x). 

Figure 2.3.1.2 Typical parametrized F-function. 
98 

Figure 2.3.1.3 Mach-plane cuts through an actual geometry, and the corresponding features in the equivalent 

area distribution. 
99 

Figure 2.3.1.4 Data from studies of boom reduction potential for airplanes of the 765-076 size and 

configuration. Boom PLdB was minimized using parametrized F-function with constraints on equivalent 

area. a) PLdB versus overall volume of the equivalent body (volume + lift), as fraction of the baseline volume. 

b) Forward and aft components of PLdB versus volume. 
100 

Figure 2.3.1.5 F-function of two best points from studies shown in Figure 2.3.1.4 

Figure 2.3.1.6 Ground signature from 50,000ft, according to MDboom, of the two best points from study 

shown in Figure 2.3.1.4. 
101 

Table 2.3.2.1. Pressure Drag variation with CG according to Tranair on Design 3. 

Figure 2.3.2.1 Contours of local Mach number from Tranair on baseline and Tranair designed 

configurations. 
102 

Figure 2.3.2.2 765-076E External Views 

OAL 154 ft

Span 86 ft

OAH 29 ft

TOGW 180 k lbs

Fan Dia 75 in
103 

Figure 2.3.2.3 765-076E Performance Summary 

Configuration

Wing Area (ESDU)

Wing Area (Gross)

Thrust per engine

Fan Diameter (in)

Engine Weight

TOGW

OEW

Payload 

Total Fuel 

Block Fuel

Range

Supersonic Cruise

Mach

Altitude

Weight

CL

L/D

SFC

Climb (pinch point)

Mach

Altitude

Weight

CL

L/D

Thrust

SFC

ROC

Bal Field Dist. (100% pwr)

W/S (takeoff)

FOM

* Range and fuel load reduced until 300 fpm ROC constraint is met * Range and fuel load reduced until 300 fpm ROC constraint is met

V1.2

Mar 16

BSE 04B Engine

2517

3344

43,000

75.4

9612

171,793

92,950

6,300

73,171

62,029

3200*

1.6

49,895

159,939

0.1457

8.513

0.9237

1.13

40,550

163,691

0.1910

8.506

19,966

0.9183

302

5309 (58%)

68.3

1.55

Mach 1.6, 30 PAX, 

variable range

Configuration

Wing Area (ESDU)

Wing Area (Gross)

Thrust per engine

Fan Diameter (in)

Engine Weight

TOGW

OEW

Payload 

Total Fuel 

Block Fuel

Range

Supersonic Cruise

Mach

Altitude

Weight

CL

L/D

SFC

Climb (pinch point)

Mach

Altitude

Weight

CL

L/D

Thrust

SFC

ROC

Bal Field Dist. (100% pwr)

W/S (takeoff)

FOM

* Range and fuel load reduced until 300 fpm ROC constraint is met * Range and fuel load reduced until 300 fpm ROC constraint is met

V1.2

Mar 16

BSE 04B Engine

2517

3344

43,000

75.4

9612

171,793

92,950

6,300

73,171

62,029

3200*

1.6

49,895

159,939

0.1457

8.513

0.9237

1.13

40,550

163,691

0.1910

8.506

19,966

0.9183

302

5309 (58%)

68.3

1.55

Mach 1.6, 30 PAX, 

variable range
104 

Figure 2.3.3.1 765-076F configuration. 
105 

Figure 2.3-2. 765-076E and -076F Geometry Comparison 
106 

Figure 2.3.3.3 Boom PLdB and ground signature for the 765-076F. 
107 

Figure 2.3.3.4 Pareto optimization starting from the 765-076E. Minimizes PLdB and maximizes L/D using 

body design variable while constraining cabin volume and overall body volume to be at least the baseline. 
108 

3.0 Propulsion Development 

3.1 Trade Space Study Engines 

The initial concept development work prior to final downselect employed proprietary Boeing Study 

Engines (BSE). They were selected by choosing a relevant airframe configuration, installing one of several 

available engines, then sizing both thrust and wing area to maximize range. Any other impact on the 

configuration was also assessed in case it might change the simple conclusion that range decides the best 

engine. 

As an example, various engines were installed on the 1080-2015 (“2015TC”) airframe and wing area and 

thrust were sized to maximize range at various Mach for 300,000lb MTOW. Figure 3.1.1 shows the 

resulting range. For this airframe, the BSE-2 engine provides superior range at Mach 1.8, and seems to be 

the natural choice. Figure 3.1.2 also supports the BSE-2 engine because wing area is minimized, 

suggesting minimum airframe weight. Also on the plot, in the gray patches at the ends of some bars, is 

evidence that wing area can vary, in concert with the engine, without greatly affecting range. But even the 

lower end of that range is not lower than the BSE-2 for all engines except the BSE-1601, and that engine 

provides significantly less range. Figure 3.1.3 shows the resulting thrust, and concludes the engine 

selection with the BSE-2 having the minimum sea level static thrust. Conventionally, that means the BSE2 engine should be lighter and cost less than the others. 

An identical exercise was performed with a different set of smaller engines installed on a smaller aft wing 

airplane at 180,000lb MTOW in an effort to choose a best engine for studying the 30 passenger low boom 

concept. 

For the engine cycle trades in Section 2.0, a series of six candidate engines cycles were developed by 

Georgia Institude of Technology (GIT), three from P&W and three from RR, were evaluated. These were 

provided to the MDAO model in the form of installed thrust and fuel flow tables for a reference thrust. 

Results from those engines have been shown in Section 2.0. 

3.2 RR NPSS Development and Cycle Trade Parameters 

Georgia Tech, with guidance from Rolls Royce, the only participant in Option 2, modeled a dual spool, 

mixed-flow turbofan (MFTF) engine with a high pressure ratio fan. The MFTF architecture was selected 

because it is a proven architecture that has been used for years, and its modeling characteristics are well 

understood. A simple, block schematic of the MFTF is shown in Figure 3.2.1 – NPSS Model Schematic. 

The engine technologies, such as component efficiency, were modeled using public domain data which was 

then modified to reflect the technology level that will be ready for entry into service in the 2025 timeframe. 

Georgia Tech iterated with Rolls Royce until both groups thought that the technology levels modeled were 

appropriate. A list of all duct pressure drops and component design efficiencies can be found in Table 3.2.1 

– MFTF Design Point Efficiencies for 072B Study (Including Tech Level Effects). The baseline 

efficiencies are not tied to a particular year, and may be adjusted by the Technology Level Effects described 

in Table 3.2.2. The Technology Level Effect used for the 072B and 076E configurations is 3.0. 

The NPSS model was created using several modularized function files that allowed for easy updates to the 

model as new data or results were made available. A brief description of the NPSS file naming convention 

is shown in Table 3.2.3 – NPSS File Naming Convention. 
109 

3.2.1 - Engine Cycle Set-up and Control Methodology 

MFTF Cycle Setup at SLS 

The MFTF is a parametric model that allows the user to select values for the cycle parameters listed in 

Table 3.2.4. These parameters were chosen because they have the largest impact on engine characteristics 

of importance to this study including specific thrust, specific fuel consumption, lapse rate, fan diameter, 

and jet velocity. The NPSS solver is used to satisfy all of the following conditions at the SLS cycle setup 

point. These are high level discussions; however more detailed descriptions of the engine model are 

presented in subsequent sections. 

• FPR is set directly from the user input. 

• OPR is set directly from the user input by varying the on design HPC pressure ratio. 

• Maximum T41 is a constraint and is used during maximum power calculations throughout 

flight envelope including takeoff conditions. 

• Throttle ratio is used to set T41 at the SLS condition. 

• Net Thrust at SLS is set by varying the SLS engine mass flow. 

• Bypass ratio is varied by the solver until the user input extraction ratio is achieved. 

• The cooling flows are sized using Equation in conjunction with Figure 3.2.2 through 

Figure 3.2.5. A fixed value of maximum T3 is used to size the cooling flows for all engine 

cycles. This was done to facilitate numerical convergence and decrease run time. 

• During cycle setup the engine design point efficiencies from Table 3.2.1 are used. Size 

effects from Figure 3.2.6 through Figure 3.2.13 are applied to the base efficiencies. 

• The engine is sized with the inlet installed.

Off-Design Operation 

In off design mode the NPSS solver is reconfigured to calculate engine operating parameters, namely thrust 

and fuel flow, at any given ambient conditions. 

• Maximum Power / Maximum Cruise / Maximum Climb 

– Vary fuel flow to run to 100% corrected fan speed unless constrained by max T4.1 limit 

or max T3 limit. 

– Exhaust nozzle throat area varied to maintain 22% Fan SMN. 

• Part Power 

– Fuel flow is varied to run to desired percentage of max power net thrust. 

– Exhaust nozzle throat is area held fixed at current maximum power value. 

• Throttle Hook 

– Function to run several points from max power to idle power to generate “hooks” at a 

specified altitude and Mach number 

• Runs Engine at maximum power. 

• Computes thrust target = max power thrust * percentage. 

• Executes model at desired part power setting. 

– The main use of this function is to generate thrust and fuel flow data for the Cases engine 

data pack. 

Inlet Sizing and Operation 

The inlet capture area is sized at a user specified top of climb altitude and Mach number. After the initial 

engine setup at SLS conditions the engine is ‘flown’ to the top of climb and run at maximum power. The 

inlet capture area is varied by the NPSS solver until the inlet size is matched to the cruise conditions using 

the recommended Mass Flow Ratio (MFR) from the inlet maps. This minimizes spillage drag and ensures 

better installed engine performance. Inlet performance is calculated during off-design using inlet recovery, 

spillage drag, bleed drag, and bypass drag maps provided by Boeing, shown in Figure 3.2.14, Figure 3.2.15, 

and Figure 3.2.16. 
110 

Nozzle Gross Thrust Coefficient 

Nozzle losses were modeled using a gross thrust coefficient, or Cfg. Boeing provided a nozzle Cfg curve 

which provides the gross thrust coefficient as a function of nozzle pressure ratio. NPSS calculates nozzle 

pressure ratio at a given flight condition; the corresponding Cfg is then applied to the gross thrust in order to 

account for losses. Figure 3.2.17 shows the variation of Cfg with nozzle pressure ratio. 

Takeoff De-rating 

In the 072B initial studies, summarized in Table 3.2.5, the MFTF could not simultaneously satisfy the 

vehicle range, fan diameter, and takeoff jet velocity requirements. Further investigation revealed that the 

flight condition that constrained the engine size was usually the top of climb point; this resulted in excess 

thrust at the takeoff condition. Since the aircraft did not need maximum available power to meet the 10,000 

ft takeoff field length requirement the engine could be de-rated for takeoff conditions. A user specified jet 

velocity input was created that limits the engine takeoff power to meet the user input. This allowed the 

MDO environment to conduct trade studies between takeoff field length and jet velocity. Jet velocity was 

chosen as a parameter because it is directly related to the jet noise requirement. The de-rating logic is as 

follows: 

• Once engine is setup it is run in off design mode at the takeoff condition. (1,000 ft / 0.255 

Mach) 

• The engine is then throttled back until the specified exit jet velocity is attained. 

• The corrected fan speed at this de-rated power is recorded as pcn2takeoff. 

• For the takeoff power flight envelope the engine is operated by varying the fuel flow until 

the corrected fan speed is equal to pcn2takeoff. 

Creating the CASES Engine Data Pack 

The CASES data pack allows for different thrust and fuel flow values to be specified for takeoff, climb, and 

cruise. Maximum cruise and climb were both defined by the maximum power definition discussed 

previously. If the engine is de-rated to a particular takeoff jet velocity then the de-rating logic is used to 

determine takeoff power available, otherwise takeoff is defined as normal maximum power operation. 

Throttle hooks are generated for each flight condition defined in the flight envelope in Figure 3.2.18. 

3.2.2 Parametric Model Effects 

Several “public domain” curves were used in order to account parametric effects as the engine cycle was 

changed. For example, the leakage cooling flows are dependent on the combustor exit temperature. 

Iteration with Rolls Royce occurred until they were satisfied that the “public domain” curves represented 

the appropriate technology levels without divulging proprietary information. The following effects are 

included in the engine model: 

• HPC Size Effect 

• HPT and LPT Size Effects 

• Technology Levels 

• Fan, HPC, HPT, and LPT Reynolds Effects 

• HPT and LPT Cooling Flows 

• Fan and HPC Loading (Effect of design point efficiency vs. design pressure ratio) 

• Weight Estimation 

HPC Size Effects 

The HPC size effect data came from a NASA report by Niedzwiecki, titled “Small Engine Technology 

Programs”. The “current technology” curve shown in Figure 3.2.19 was selected from the report and the 
111 

only adjustment was to convert the polytropic efficiency to a delta on adiabatic efficiency. This was 

necessary because NPSS uses adiabatic efficiency in map and performance calculations. 

Turbine Size Effects 

The HPT and LPT size effect data came from a NASA report by Niedzwiecki, titled ““Small Engine 

Technology Programs”. The ‘current technology’ curve shown in Figure 3.2.20 the corrected flow axis to 

represent 40 lbm/s and 10 lbm/s, respectively. 

Technology Levels 

The technology levels are adders relative to a fixed baseline value. By changing the overall technology 

level of the engine all of the components are adjusted simultaneously. 

Reynolds Effects 

Reynolds effects were based on Bullock’s “Analysis of Re and Scale Effects on Performance of 

Turbomachinery” (J. Engineering for Power, July 1964, pp.247-256). Equation describes the Reynolds 

effect calculation implemented in the MFTF. The NPSS model also has relationships to relate the flow and 

pressure ratio scalars to the efficiency adder. These are shown in the equations below. Graphs of the 

resulting Reynolds effects for each component were shown in Figure 3.2.6 through Figure 3.2.13. The 

NPSS compressor and turbine elements and Reynolds number sockets were modified to accept these 

curves. 

0.2

1.0 1

1 −

=

= −

− RNI

η RNI

η

Equation 3.2.1 MFTF Reynolds Effects Equation 

P

P

s s w

s w s eff

_ Pr _

_ _ 1

=

Δ = = +

η

η

Equation 3.2.2 MFTF Reynolds Effects on Pressure Ratio and Flow Rate 

Turbine Cooling 

A Coolit type algorithm is used with a cooling effectiveness factor of 1.0. This is equivalent to full 

coverage film cooling. The cooling logic in the MFTF takes the single blade row cooling calculations from 

Coolit and adapts them to a “Total Chargeable Cooling” curve. This assumes the chargeable cooling curves 

apply to the HPT first row blade and that the HPT first row blade accounts for 30% of the total turbine 

cooling. A metal temperature of 2000oF is used to calculate the required cooling. This metal temperature 

was chosen because it represents the compromise between an aggressive temperature for a nickel based 

alloy, and a conservative temperature for a CMC material. 

( )

T ( ) T CCA T

T Deterioration T

Max Max

Max metal

− − Δ

+ Δ − = _ 4.1, 3,

4.1, φ

Equation 3.2.3 Cooling Effectiveness 

The curve for total chargeable cooling is shown in Figure 3.2.2. Cooling effectiveness is defined in 

Equation above. 

Total Leakage is defined as a linear function of the Turbine Rotor Inlet Temperature and is shown in Figure 

3.2.3. 

Total cooling flow, shown in Figure 3.2.4, was distributed across the HPT and LPT to provide reasonable 

levels of cooling for the temperatures being investigated. The fractions in Figure 3.2.4 are multiplied by the 

total cooling air fraction obtained from Figure 3.2.2 to obtain the cooling flow for each component. The 
112 

fractions listed in Figure 3.2.5 represent the corresponding leakage flow distribution and are multiplied by 

the total leakage obtained from Figure 3.2.3 to set the leakage flow. 

Fan and Compressor Loading 

The fan and HPC polytropic efficiencies are parametric functions of the fan and compressor loading. A 

loading curve from Hill & Peterson, Mechanics & Thermodynamics of Propulsion (Addison-Wesley, 2nd

edition, 1992) was supplemented with data from Creason & Baghdadi, Design and Test of a Low Aspect 

Ratio Fan Stage, AIAA paper 88-2816. This provided the fan and HPC loading curves shown in Figure 

3.2.21. Since the Fan Pressure Ratio (FPR) varied from 1.8 to 3.0, the multistage fan loading curve was 

used. 

Weight Estimation 

Weight estimates are based on a simplified methodology provided by Rolls Royce and described by the 

equations below. The equations use basic flow scaling of components and Georgia Tech modified the 

reference values based on previous experiments with WATE++ studies. The reference values used in the 

various weight equations are found in Table 3.2.6. Specific works per stage values came from matching 

known cycles to stage counts for various engines. Since these values came from current engines they may 

be somewhat conservative for the N+2 timeframe. 

Equation 3.2.4 MFTF Weight Estimation Equation 

1.1 n

Nstg_Ref

Nstg

Wc_Ref

Wc

Weight_Ref

Weight ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

∗ ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ =

Equation 3.2.5 Turbomachinery Weight Equation 

( ) SpecificWork/Stage

SpecifcWork Nstg =

Equation 3.2.6 Stage Count Equation 

1.1

Wc_Ref

HPC_Exit_W c

Weight_Ref

CDN_Weight ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ =

Equation 3.2.7 CDN Weight Equation 

1.1

Wc_Ref

HPC_Inlet_ Wc

Weight_Ref

AGB_Weight ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ =

Equation 3.2.8 AGB Weight Equation 

1.1

Wc_Ref

HPC_Inlet_ Wc

Weight_Ref

Misc_Weigh t

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ =

Equation 3.2.9 Misc Weight Equation 

3.2.3 - MFTF Configuration and Cycle Selection for the 072B 

Georgia Tech investigated the design space shown in Table 3.2.4 to find three candidate engine cycles for 

the 072B. From initial design space explorations it was found that the range of cycle parameters shown in 

Table 3.2.4 provides good performance for the 072B configuration. It was not possible to create an engine 

cycle with the assumed technology levels that simultaneously satisfied the jet velocity, fan diameter, 
113 

mission fuel burn requirements, and vehicle thrust requirements. For this reason Georgia Tech provided 

three engine cycles for mission analysis, one which met the jet velocity constraint, one which came closer 

to the fan diameter constraint, and one which was a compromise between the two objectives. This allowed 

Boeing to perform the trades required to determine which cycle was the best selection for the 072B. These 

cycles are not optimum solutions since they are a result of a discrete search of the design space; however, 

they are still useful in performing trade studies and may be good candidates as initial conditions in a 

gradient based search. 

For the design space exploration 10,000 engine cycles were run using a Latin Hypercube which is a type of 

Design of Experiments, or DoE. By using a DoE the design space can be intelligently explored with far 

fewer engine cycles than would be required with a random grid search. A multivariate plot of the complete 

design space with all cycle inputs and relevant outputs such as fan diameter, SFC, and jet velocity is shown 

in Figure 3.2.22. 

Cycles that did not meet the thrust requirements of the 072B at various flight conditions were filtered out of 

the design space, shown in Figure 3.2.23. The absolute minimum jet velocity of the cycles that survived the 

thrust requirement filtering was 1258 ft/s which did not meet the requirement of 1100 ft/s. The minimum 

fan diameter remaining was 86 inches which did satisfy the constraint imposed by Boeing for the 072B. 

In order to select the three cycles to provide to Boeing, Georgia Tech examined the fan diameter and jet 

velocity Pareto front shown in Figure 3.2.24. Three cycles were selected, the ones with absolute minimum 

fan diameter, absolute minimum jet velocity, and a compromise solution. In Figure 3.2.24 these cycles are 

represented with the labels Case A, Case B, and Case C respectively. The compromise solution was 

selected using TOPSIS, or Technique for Order Preference by Similarity to Ideal Solution. By weighting 

the relevant importance of jet velocity and fan diameter a TOPSIS finds the cycle with the closest Euclidian 

distance to the ideal solutions of 1100 ft/s and 86”. The three selected cycles and their performance is 

summarized in Table 3.2.5 and Table 3.2.7. As is to be expected the low fan diameter cycle has the highest 

FPR and a correspondingly high jet velocity. Also of interest is that the compromise cycle, Case C in 

Figure 3.2.3.3, has a lower maximum T4.1 than the other two selected cycles. This is because these cycles 

resulted from a discrete search of the design space. There are other cycles that are very close to Case C on 

the Pareto front that have different maximum T4.1’s. This reiterates that need to use these cycles as initial 

points in a further MDO optimization. 

3.2.4 - MFTF Updates for 076E Configuration 

Minor updates to the MFTF NPSS model were made for the study of the low-boom 076E configuration. 

After discussing the results of the 072B study internally and with feedback from Rolls Royce, Georgia 

Tech reached the conclusion that the technology levels used in the 072B version of the MFTF were too 

conservative. The base efficiencies of the turbomachinery components and the burner were increased. The 

new efficiencies for the engines used in the 076E study are shown in Table 3.2.8. The efficiencies reflect 

the baseline efficiency added to the technology effect previously discussed in Section 3.2.2 Parametric 

Model Effects. 

3.3 The coupled NPSS/MDA Tool 

The integration of NPSS into the MDAO process allowed the capability to optimize engine-airframe 

combination in a way previously unavailable. Without NPSS, the ModelCenter model used assumed 

nacelle scaling correlations to grow and shrink the nacelle based on the thrust required, while the thrust and 

fuel flow for the engine was merely scaled within CASES. NPSS allowed an engine deck to be created 

with each iteration, providing a more accurate sized engine for different thrust requirements. As the 

reference thrust for the engine deck is the required thrust, there is no longer any need to scale thrust and 

fuel flow in CASES, increasing the accuracy of the analysis. In addition, NPSS estimates the engine 

diameter, which allows more accurate inputs to geometry than the assumed correlations. 
114 

For the 765-072B configuration, the nacelle geometry was allowed to change with changes to the engine 

parameters, which provided a wide design space for the parametric engine. A weight was estimated based 

on an engine thrust-to-weight assumption of between 4.25 and 5.0. 

When the 765-076E configuration was analyzed, an addition to NPSS was made available that estimated 

the engine weight internally, removing the necessity to roughly estimate engine weight based on thrust-toweight. The 765-076E outer-mold-lines were fixed because of the low-boom requirements, and so the 

nacelle diameter was not allowed to grow, though it was allowed to shrink. 

The first integration was between NPSS and the -072B species MDAO model. To check out this 

integration, we chose to do a comparison between the vehicle performance from the GTRR Compromise 

engine cycle dataset provided for the downselect trade studies (Section 2.0) and the NPSS output 

simulating the cycle inputs for this engine. On any design study, things are constantly changing, so it was 

impossible to do a completely “apples to apples” comparison. For example, the mission calculations had 

been revised to incorporate a maximum cruise altitude constraint of 55,000 ft (lower than optimum), which 

came out of the Market Study. The NPSS model had also been updated with lessons learned from the midterm review, so the same inputs would not exactly produce the same outputs. Finally, the -072B MDAO 

aerodynamics analysis had been updated slightly as a result of completing the model validation exercise. 

Figure 3.3.1 shows a comparison of the (updated) vehicle characteristics with the GTRR compromise 

engine (Column 1) to the NPSS generated engine with the same cycle inputs (Column 2). In both examples 

the thrust and wing area had been resized to meet the ROC and W/S constraints. No attempt was made 

here to derate the engine for takeoff. So the balanced field distances shown are for an engine at 100% 

power. The conclusion is that with the revised NPSS deck integrated into the MDAO model requires about 

5% more thrust for the same cycle parameters. This additional thrust results in a slightly larger engine, 

higher TOGW, more fuel and larger wing to meet the same aircraft performance requirements. The 

resulting FOM drops from 3.44 to 3.06. 

3.4 072B Optimization 

For the cycle optimization with the -072B species MDAO model, the primary objective was to maximize 

FOM. This optimization would be subject to several constraints, as follows. 

• ROC greater than 300 fpm anywhere in the mission 

• Sufficient fuel volume in the wing (desired) or no more than 47,200 lbs of fuel in the fuselage, 

below the passenger compartment (required) 

• Takeoff W/S less than 78 psf (approximates a 155 kt approach speed) 

• Balanced field distance less than 10,000 ft 

• Takeoff Vj less than 1100 fps (derated) 

Initial attempts at the optimization showed that the design space did not have any feasible solutions for a 

10,000 ft balanced field distance with 1100 fps Vj. Therefore, Vj was changed from a constraint to an 

optimization goal. Thus we now had a multiple objective optimization problem. 

The selected optimizer for this study was Design Explorer. This is a Boeing developed optimizer, which 

has been licensed to Phoenix Integration for use in Model Center. The optimizer uses an algorithm 

(SEQOPT) that is more effective on “noisy” data or design spaces with multiple peaks and valleys than a 

traditional gradient search algorithm. Design Explorer has a limitation of a single objective function. 

Since we had a multi-objective optimization, we elected to conduct the following three optimizations and 

compare the results. 

• Maximize FOM 

• Minimize Vj 

• Maximize FOM/Vj 

We anticipated that each of these would result in a significantly different “optimum” engine cycle. 
115 

The design variables for the optimization and their allowable ranges are as follows. 

• Lower Upper

• Fan Pressure Ratio (Fan_PR) 1.8 2.9 

• Overall Pressure Ratio (Overall_PR) 25 35 

• Extraction Ratio 0.9 1.1 

• Max Turbine Temperature (Max_T41) (deg R) 2700 3700 

• Throttle Ratio 1 1.2 

• Inlet Design Altitude (ft) 40,000 55,000 

• Jet Velocity (Vj) (fps) 1100 2000 

• Net Thrust (lbs) 50,000 70,000 

• Wing Area (sgross_ft) 4000 6000 

The limits for the cycle design variables were established by Georgia Tech and Rolls Royce as 

representative of N+2 technology level. The limits for thrust and wing area were chosen based on 

experience from the earlier trade studies. The upper limit for the inlet design altitude was the same as the 

max cruise altitude constraint. We observed that the inlet design altitude always optimized at the upper 

limit of 55,000 ft. It was eventually eliminated as a design variable and simply set at 55,000 ft. This 

reduced the dimensionality of the optimization problem. The lower limit for the jet velocity was the project 

goal for takeoff. The upper limit was made as low as possible, but still result is feasible design solutions 

for the 10,000 ft balanced field distance. The reader will note that Vj is both a design variable and an 

optimization objective. There are actually two Vj design variables in the NPSS model (zVjet, Vj). One is 

an input and the other is an output. When the NPSS derate switch is turned on, these two variables are 

made equal. For the maximize FOM the derate switch was turned off. For the other two optimizations, the 

derate switch was turned on. 

The optimization needs a set of starting values for the design variables. It seems to work better if the set 

results in a feasible design. The values for the GTRR Compromise engine from the trade study analysis 

were used as the starting point. This concept does not exactly match the concept shown in Figure 3.3.1 

because the inlet design point was changed from Mach 1.8, 43000 ft to Mach 1.6, 55,000 ft for this 

comparison. This concept was designated the Benchmark. Each of the optimizations was compared to it to 

determine if it improved FOM, Vj, or both. For design space with many peaks and valley, it is possible that 

a different starting point for the design point for the design variables will result in a different optimum 

solution. This is known as a local optimum, as opposed for the global optimum that we hoped to find. In a 

large, multi-dimensional design space is very difficult to determine for sure if an optimization has found a 

true global optimum. For this study, once an optimum was found, we tried a couple of different starting 

points to see if they would arrive at the same optimum solution. This is clearly not a rigorous way to make 

the determination, but it provided some level of confidence that a global optimum was obtained. 

The results of the first optimization to maximize the FOM are shown in Figure 3.4.1, along with the 

Benchmark GTRR compromise engine cycle. The optimization changed all the design variables with a 

significant increase in fan pressure ratio and Max T41. This allowed the thrust and wing area to be 

reduced, while still meeting the constraints. The FOM for the optimum concept increased from 3.14 to 

3.43. However, the Vj also increased from 1375 to 1640 fps for a 10,000 ft balanced field distance with a 

derated engine. 

The results of the second optimization to minimize Vj are shown in Figure 3.4.2 along with the Benchmark. 

The optimization again changed all of the design variables. The max T41 was again significantly 

increased. This time, however, the fan and overall pressure ratios were reduced. The required thrust 

increased, but the wing area decreased. The optimized FOM increased slightly from 3.14 to 3.17 and the 

optimized Vj significantly dropped from 1375 to 1194 fps for a 10,000 ft balanced field distance with a 

derated engine. The Vj did not achieve the goal of 1100 fps. 

The results of the third optimization to maximize FOM/Vj are shown in Figure 3.4.3 along with the 

Benchmark. The optimization again changed all of the design variables. The max T41 went to the limit of 

3700 deg R. The overall pressure ratio increased, but the fan pressure ratio decreased. The required thrust 
116 

increased, but the wing area decreased. The optimized FOM increased from 3.14 to 3.22 and the Vj 

dropped from 1375 to 1204 fps. 

The three optimizations are shown compared to each other in Figure 3.4.4. In deciding how to choose from 

the three optimizations for the competing objectives, we observed that all resulted in an FOM greater than 

the project goal of 3. None resulted in a Vj below the project goal of 1100 fps. Therefore, we selected the 

optimization that provided the lowest Vj (optimization 2 – minimize Vj). 

Appendix 1 contains some visual details of the 072B. 

3.5 076E Optimization 

As discussed in section 2.3.4, the -076E species MDAO model had fixed external mold lines and a 

maximum allowable takeoff gross weight, due to the boom signature constraints. This meant that the 

optimization objectives, constraints, and design variables would be different for the -076E. The -072B 

optimization used a fixed 4000 nm range and variable TOGW, while the -076E has a fixed 180,000 lb 

maximum TOGW and a variable range. We elected to conduct the following three optimization for the -

076E and compare the results. 

• Maximize Range 

• Minimize Vj 

• Maximize Range/Vj 

. This optimization would be subject to several constraints, as follows. 

• ROC greater than 300 fpm anywhere in the mission 

• Sufficient fuel volume in the wing 

• Takeoff W/S less than 78 psf (approximates a 155 kt approach speed) 

• Balanced field distance less than 10,000 ft 

• Fan diameter less than 75.4 inches 

• TOGW less than 180,000 lbs 

 

Many of these are the same constraints as for the -072B. The fan diameter constraint was added to reflect 

the fixed external mold lines of the -076E. The design variables for the optimization and their allowable 

ranges are as follows. 

• Lower Upper

• Fan Pressure Ratio (Fan_PR) 1.9 3.4 

• Overall Pressure Ratio (Overall_PR) 20 50 

• Extraction Ratio 0.9 1.1 

• Max Turbine Temperature (Max_T41) (deg R) 2800 4000 

• Throttle Ratio 1 1.2 

• Jet Velocity (Vj) (fps) 1100 1300 

• Net Thrust (lbs) 30,000 50,000 

• Range (nm) 3500 4000 

The limits for the cycle design variables were extended beyond what was used for the -072B by Georgia 

Tech and Rolls Royce after preliminary optimizations showed it would be difficult to find feasible solutions 

for the -076E. These limits are representative of something between N+2 and N+3 technology levels. The 

4000 nm upper limit for the range was the recommended value from the earlier trade studies, while the 

3500 nm lower limit was the least we could tolerate with a transcontinental mission. The thrust limits were 

what preliminary analyses showed would result in feasible designs. The 1100 fps lower limit for Vj was 

the project goal, while the 1300 fps upper limit was what preliminary analyses showed would result in 

feasible designs. The philosophy was that even while maximizing FOM, we didn’t want the Vj to get too 

high. The wing area was eliminated as a design variable, due to the fixed external mold lines. The range is 

both a design variable and an optimization objective. In the MDAO model there is an input range and an 
117 

output range. The first is the design variable and the second is the objective. When the MDAO model 

iterates to close the mission solution, these two variables are made equal. 

The starting point cycle for the -076E optimization was selected to be the optimum cycle from the -072B 

optimization the resulting concept is designated the Benchmark. The point of departure engine cycle for 

the -076E (Section 2.5) was the proprietary BSE 04B engine. It was not possible to simulate this cycle in 

NPSS for this non-proprietary study, so it was not used as the benchmark. 

The results of the first optimization to maximize the FOM are shown in Figure 3.5.1, along with the 

Benchmark GTRR compromise engine cycle. For this optimal solution, the Max T41 went to its upper 

limit of 4000 deg R. The overall pressure ratio also increased, but the fan pressure ratio decreased. To 

meet the ROC constraint the thrust increased. The range increased from 3700 nm to 3910 nm. This is 

reasonably close to the goal of 4000 nm. The Vj decreased from 1510 to 1250 fps for a 10,000 ft balanced 

field distance with a derated engine. 

The results of the second optimization to minimize Vj are shown in Figure 3.5.2 and yielded a significantly 

different cycle. Both the fan and overall pressure ratios decreased from the bench mark. The max T41 

increased, but did not reach its limit. The range went to its lower limit of 3500 nm in an attempt to 

minimize the TOGW and thrust. The resulting Vj decreased from 1510 to 1164 fps, which is close to the 

project goal of 1100 fps. 

The results of the third optimization to maximize range/Vj are shown in Figure 3.5.3 and again yielded a 

significantly different cycle. The max T41 again went to the upper limit of 4000 deg R. The overall 

pressure ratio increased, but the fan pressure ratio decreased. The thrust also increased. In this 

optimization, the range increased slightly from 3700 to 3799 nm and the Vj decreased significantly from 

1510 to 1197 fps. 

The three optimization results are compared in figure 3.5.4. There were two competing project objectives: 

4000 nm range and 1100 fps Vj. None of the three optimizations achieved either of these objectives. The 

max range optimization achieved the best range, the min Vj optimization achieved the lowest Vj, and the 

max range/Vj optimization was a compromise. Without any clear knowledge of the relative importance of 

range versus Vj, we chose the compromise optimization 3 (max range/Vj) for the preferred engine cycle. 

3.6 Selected Cycle Parameters 

The selected optimum engine cycle parameters for the -072B and -076E concepts are as follows. 

• 765-072B 

o Fan Pressure Ratio (Fan_PR) 2.10723 

o Overall Pressure Ratio (Overall_PR) 34.4961 

o Extraction Ratio 1.05938 

o Max Turbine Temperature (Max_T41) (deg R) 3666.8 

o Throttle Ratio 1.12539 

o Inlet Design Altitude (ft) 55,000 

o Derate Jet Velocity (Vj) (fps) 1194 

o Net Thrust (lbs) 64,688 

• 765-076E 

o Fan Pressure Ratio (Fan_PR) 2.20176 

o Overall Pressure Ratio (Overall_PR) 39.2188 

o Extraction Ratio 0.9 

o Max Turbine Temperature (Max_T41) (deg R) 4000 

o Throttle Ratio 1.10039 

o Inlet Design Altitude (ft) 55,000 

o Derate Jet Velocity (Vj) (fps) 1197 

o Net Thrust (lbs) 41,133 
118 

Both the -072B and -076E are twin engine concepts. The -072B is a larger 100 passenger aircraft, while 

the -076E is a smaller 30 passenger aircraft. As expected, the -072B engine produces more thrust than the -

076E engine. The -076E is a more constrained aircraft which required higher levels of Max T41, overall 

and fan pressure ratios in order to achieve reasonable flight performance. 

The cycle selected by Boeing for the 076E configuration has a very high turbine inlet temperature of 

4000oR. For the 076E study Georgia Tech thought that by examining higher TIT’s the resulting increased 

engine efficiency might help to bridge the gap between the vehicle performance and requirements. This 

high temperature takes advantage of future material technologies that will be present at the time of entry 

into service. The cooling flow model in NPSS assumes a metal temperature of 2460oR. As discussed in 

3.2.2 Parametric Model Effects, this temperature, along with the appropriate cooling flows, is compatible 

with the material available in the N+2 timeframe. Due to the increased allowable metal temperature the 

total cooling and leakage flow for the selected cycle are at a reasonable value of 32.2% of the HPC inlet 

flow. Additionally the difference between the cooling gas temperature and gas path temperature is similar 

to current engines. 

The MDO optimization chose a low FPR because it lowers the jet velocity by increasing the engine bypass 

ratio. The secondary effect of low FPR is a worse thrust lapse rate which means that the engine has to be 

even more oversized at takeoff to meet top of climb thrust requirements. Therefore the engine must be 

further de-rated to lower the jet velocity. 

The throttle ratio corresponds to a theta break of approximately Mach 1.55 at the cruise altitudes. This is to 

be expected since the engine should be operating near the maximum T3 and T4.1 in order to achieve the 

best cycle efficiency. 

The amount of cooling air required for the high TIT of 4000R may be reduced by the addition of cooled 

cooling air (CCA) technology. In the NPSS model provided to Boeing for the 076E study the CCA feature 

was not enabled; however the CCA feature is enabled in the final model delivered to Boeing. The CCA 

implementation, as shown in Figure 3.5.5, cools compressor discharge air using fan bypass duct air before 

it is used to cool the turbine blade. The reduction in cooling air required is accounted for by the CCA_ ΔT 

term in Equation 3.2.3. 

To draw the 2-D flowpath of the down-selected cycle Georgia Tech elected to use NASA’s WATE++ tool 

due to its integration with NPSS. Georgia Tech designed a flowpath using WATE++ and iterated with 

Rolls Royce until they were satisfied with the final WATE++ drawing. Rolls Royce used the final 

WATE++ output to draw the final CAD model. 

 
119 

Tables and Figures for Section 3.0 

Figure 3.1.1 Thrust and area sizing results for various engines on a scaled 2015TC airframe. 

0

1000

2000

3000

4000

5000

6000

7000

BSE 1 BSE 2 BSE 1601 3670.51.27 3870.47.27

ultimate MFTF

3870.47.27

Mach 2 comb

3870.47.27

Mach 2 ultimate

Engine

Range - nmi

Cruise Mach 1.6

Cruise Mach 1.8

Cruise Mach 2.0
120 

Figure 3.1.2. Wing area resulting from engine/airframe sizing on a scaled 2015TC airframe. 
121 

Figure 3.1.3. Thrust resulting from engine/airframe sizing on a scaled 2015TC airframe. 

0

10000

20000

30000

40000

50000

60000

70000

BSE 1 BSE 2 BSE 1601 3670.51.27 3870.47.27

ultimate MFTF

3870.47.27

Mach 2 comb

3870.47.27

Mach 2 ultimate

Engine

Sea Level Static Thrust - lb

Cruise Mach 1.6

Cruise Mach 1.8

Cruise Mach 2.0
122 

Table 3.2.1 MFTF Design Point Efficiencies for 072B Study (Including Tech Level Effects) 

NPSS Component Name 

(Description)

Value

Duct1 (Swan Neck Duct) Pressure loss = 1.0%

Fan design point efficiency 0.90 (polytropic)

Duct2 (HPC to Burner) pressure loss = 2.0%

HPC design point efficiency 0.886 (polytropic)

Fuel LHV 18400

Burner efficiency 0.997

Burner design point pressure loss = 2.7%

HPT design point efficiency 0.917 (adiabatic)

LPT design point efficiency 0.93 (adiabatic)

Duct3 (TEGV duct) pressure loss = 1.0%

Duct5 (Bypass duct) pressure loss = 5.0%

Duct4 (Tailpipe) pressure loss = 0.5%

Nozzle Discharge flow coefficient and velocity 

coefficient are functions of NPR. Curves 

were provided by Boeing.

Table 3.2.2 Technology Level Effects 

Tech. Level Fan ηpoly HPC ηpoly Burner η HPT ηad LPT ηad

0 0 0 0 0 0

1 0.008 0.012 0.002 0.01 0.01

2 0.016 0.025 0.005 0.03 0.02

3 

(Study Value) 0.025 0.030 0.007 0.05 0.03

4 0.035 0.035 0.009 0.06 0.04
123 

Table 3.2.3 NPSS File Naming Convention 

GT MFTF NPSS Naming Convention

.mdl file Defines the engine model general configuration

.fnc files Declares variables and defines functions used in the .run file

.run file Executes the design and off-design points

.map files Defines operating characteristics of the engine components

.view files Defines the output files dynamically

.case_xxxx Defines a set of flight conditions to execute the engine model.

Table 3.2.4 Rolls Royce MFTF Design Space for 072B 

Rolls Royce MFTF Design Space

Overall Pressure Ratio 26 – 36

Fan Pressure Ratio 1.8 – 3.0

Takeoff Thrust 68,000 –90,000 lbf

Extraction Ratio 0.9 – 1.1

T41 max 3000 – 3800 °R

Throttle Ratio 1.0 – 1.16
124 

Table 3.2.5 Summary of Cycles Selected for MFTF/072B 

Table 3.2.6 Weight Estimation Reference Values 

Component Weight_Ref Wc_Ref Nstg_Ref SpecWork/Stg n (Equation )

Fan 0.82 1.0 1 19.56 0.9

HPC 4.38 1.0 9 20.79 1.0

HPT 32.69 1.0 2 208.34 0.5

LPT 10.68 1.0 3 53.09 0.9

CDN 20.67 1.0 - - -

AGB 0.69 - - - -

Misc 4.81 - - - -
125 

Table 3.2.7 Continued Summary of MFTF/072B Cycles 

Table 3.2.8 Efficiency Updates to MFTF for 076E (Including Tech Level Effects) 

Fan 0.905 (polytropic) 

HPC 0.915 (polytropic)

HPT 0.92 (adiabatic)

LPT 0.93 (adiabatic)

Burner 0.997
126 

Figure 3.2.1 NPSS Model Schematic 
127 

Figure 3.2.2 Total Chargeable Cooling 

Figure 3.2.3 Turbine Total Leakage Flow 

Figure 3.2.4 Turbine Cooling Flow Distribution 

 

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9

Cooling Effectiveness

Total Chargeable Cooling Flow, fraction of W25 0.00

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

1500 2000 2500 3000 3500 4000

Turbine Rotor Inlet Temperature, deg F

Total Leakage Flow, fraction of W25
128 

Figure 3.2.5 Turbine Chargeable Leakage Flow Distribution 

Figure 3.2.6 Fan Reynolds Effects (Efficiency) 

Figure 3.2.7 Fan Reynolds Effects (Flow & Pressure Ratio) 

-0.030

-0.025

-0.020

-0.015

-0.010

-0.005

0.000

0.0 0.2 0.4 0.6 0.8 1.0 1.2

RNI

d_epoly 0.970

0.975

0.980

0.985

0.990

0.995

1.000

1.005

0.0 0.2 0.4 0.6 0.8 1.0 1.2

RNI

s_fl & s_pr
129 

Figure 3.2.8 HPC Reynolds Effect (Efficiency) 

Figure 3.2.9 HPC Reynolds Effect (Flow & Pressure Ratio) 

Figure 3.2.10 HPT Reynolds Effect (Efficiency) 

-0.080

-0.070

-0.060

-0.050

-0.040

-0.030

-0.020

-0.010

0.000

0.0 0.2 0.4 0.6 0.8 1.0 1.2

RNI

d_epoly 0.93

0.94

0.95

0.96

0.97

0.98

0.99

1.00

1.01

0.0 0.2 0.4 0.6 0.8 1.0 1.2

RNI

s_fl & s_pr -0.025

-0.020

-0.015

-0.010

-0.005

0.000

0.0 0.2 0.4 0.6 0.8 1.0 1.2

RNI

d_ea
130 

Figure 3.2.11 HPT Reynolds Effects (Flow & Pressure Ratio) 

Figure 3.2.12 LPT Reynolds Effects (Efficiency) 

Figure 3.2.13 LPT Reynolds Effects (Flow & Pressure Ratio) 

0.975

0.980

0.985

0.990

0.995

1.000

1.005

0.0 0.2 0.4 0.6 0.8 1.0 1.2

RNI

s_fl -0.045

-0.040

-0.035

-0.030

-0.025

-0.020

-0.015

-0.010

-0.005

0.000

0.0 0.2 0.4 0.6 0.8 1.0 1.2

RNI

d_ea 0.950

0.955

0.960

0.965

0.970

0.975

0.980

0.985

0.990

0.995

1.000

1.005

0.0 0.2 0.4 0.6 0.8 1.0 1.2

RNI

s_fl
131 

Figure 3.2.14 Inlet Recovery 

Figure 3.2.15 Inlet Spillage Drag Coefficient 
132 

Figure 3.2.16 Inlet Bypass Drag 

Figure 3.2.17 Nozzle Gross Thrust Coefficient 

Max Continuous / Max Climb / Cruise Powerlines / Idle

Altitude (ft) Mach Number

0 0 0.2 0.3 0.4 0.5

1500 0 0.2 0.3 0.4 0.5 0.6

5000 0 0.2 0.3 0.4 0.5 0.6 0.7

10000 0.3 0.4 0.5 0.6 0.7 0.8

15000 0.4 0.5 0.6 0.7 0.8 0.9

20000 0.5 0.6 0.7 0.8 0.9 0.95 0.98 1

25000 0.6 0.7 0.8 0.9 0.95 0.98 1

30000 0.7 0.8 0.9 0.95 0.98 1 1.05 1.1 1.2 1.3

33000 0.7 0.8 0.9 0.95 0.98 1 1.05 1.1 1.2 1.3 1.4

36089 0.7 0.8 0.9 0.95 0.98 1 1.05 1.1 1.2 1.3 1.4 1.5 1.6 1.8 2.0

39000 0.8 0.9 0.95 0.98 1 1.05 1.1 1.2 1.3 1.4 1.5 1.6 1.8 2.0

43000 0.8 0.9 0.95 0.98 1 1.05 1.1 1.2 1.3 1.4 1.5 1.6 1.8 2.0

47000 0.95 0.98 1 1.05 1.1 1.2 1.3 1.4 1.5 1.6 1.8 2.0

50000 0.95 0.98 1 1.05 1.1 1.2 1.3 1.4 1.5 1.6 1.8 2.0

53000 0.95 0.98 1 1.05 1.1 1.2 1.3 1.4 1.5 1.6 1.8 2.0

55000 1 1.05 1.1 1.2 1.3 1.4 1.5 1.6 1.8 2.0

60000 1.05 1.1 1.2 1.3 1.4 1.5 1.6 1.8 2.0

Thrust MCT MCL 90% 85% 80% 70% 60% 40% 20% Idle 

Figure 3.2.18 MFTF Flight Envelope 
133 

Figure 3.2.19 Public Domain HPC Size Effect 

Figure 3.2.20 Public Domain HPT/LPT Size Effect 

-0.3

-0.25

-0.2

-0.15

-0.1

-0.05

0

0 2 4 6 8 10 12

Delta polytropic efficiency

Compressor Exit Corrected Flow (lbm/s)

-0.07

-0.06

-0.05

-0.04

-0.03

-0.02

-0.01

0

0 2 4 6 8 10 12

Delta adiabatic efficiency

Turbine Inlet Corrected Flow (lbm/s)
134 

Figure 3.2.21 MFTF Fan and HPC Loading vs. Efficiency 

0.72

0.77

0.82

0.87

0.92

1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6

First Stage Pressure Ratio

Polytropic Efficiency

HPC

Multistage Fan

Single Stage Fan

LPC
135 

Figure 3.2.22 Unfiltered Rolls Royce MFTF 072B Design Space 
136 

Figure 3.2.23 Rolls Royce Filtered Design Space for 072B 
137 

Figure 3.2.24 Rolls Royce Fan Diameter/Vj Pareto Front 
138 

Figure 3.3.1 765-072B Performance Comparison for NPSS Engine Cycle 

ConfigurationWing Area (ESDU)Wing Area (Gross)Thrust per engineEngine T/WTOGWOEWPayload Total Fuel Block FuelRangeSupersonic Cruise

MachAltitudeWeightCLL/DSFC

Climb (pinch point)

MachAltitudeWeightCLL/DThrustSFCROC

Bal Field Dist. (100% pwr)W/S (takeoff)FOM

V2.5+NPSS (resized)

Mar 12

Sim RR Comp 

40995000

61,000

5

317,499146,420

21,000

150,561130,557

4000

1.6

51,731

293,649

0.1794

9.2141.012

1.13

40,550

303,826

0.2177

9.418

33,6660.8318

318

6192 (??%)

77.53.06

Note 1: V2.5 of the MDAO model has corrections to aero and engine scalingNote 2: NPSS produces different thrust & fuel flow for the same cycle inputsNote 3: None of the engines were de-rated to meet 1100 fps takeoff VjNote 4: V2.5 cases were constrained to 55,000 ft max cruise altitude

V2.5 (resized)

Mar 12

Compromise 2

37714600

58,000

5

291,098135,320

21,000

135,130116,346

4000

1.6

52,412

267,223

0.1833

9.486

1.0107

1.13

40,550

277,651

0.2162

9.797

29,6410.8533

322

7088

77.23.44

Mach 1.6, 4000 nm, 100 

ConfigurationWing Area (ESDU)Wing Area (Gross)Thrust per engineEngine T/WTOGWOEWPayload Total Fuel Block FuelRangeSupersonic Cruise

MachAltitudeWeightCLL/DSFC

Climb (pinch point)

MachAltitudeWeightCLL/DThrustSFCROC

Bal Field Dist. (100% pwr)W/S (takeoff)FOM

V2.5+NPSS (resized)

Mar 12

Sim RR Comp 

40995000

61,000

5

317,499146,420

21,000

150,561130,557

4000

1.6

51,731

293,649

0.1794

9.2141.012

1.13

40,550

303,826

0.2177

9.418

33,6660.8318

318

6192 (??%)

77.53.06

Note 1: V2.5 of the MDAO model has corrections to aero and engine scalingNote 2: NPSS produces different thrust & fuel flow for the same cycle inputsNote 3: None of the engines were de-rated to meet 1100 fps takeoff VjNote 4: V2.5 cases were constrained to 55,000 ft max cruise altitude

V2.5 (resized)

Mar 12

Compromise 2

37714600

58,000

5

291,098135,320

21,000

135,130116,346

4000

1.6

52,412

267,223

0.1833

9.486

1.0107

1.13

40,550

277,651

0.2162

9.797

29,6410.8533

322

7088

77.23.44

Mach 1.6, 4000 nm, 100 
139 

Figure 3.4.1 765-072B Optimization #1 Summary (Maximize FOM) 

• Both Optimization 1 and GTRR Compromise were resized to meet constraints • Both meet 55,000 ft “threshold” max cruise altitude limit

– Changed GTRR Compromise engine design cruise point to Mach 1.6, 55,000 ft (originally M1.8, 4

GTRR Compromise Engine

Fan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Vj=1375 fps

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Vj=1375 fps

Optimization improved on GTRR Compromise cycle at the expense of V

Optimization improved on GTRR Compromise cycle at the expense of V

Optimized Design (307)

Fan_PR=2.891406Overall_PR=36Extraction_Ratio=0.9984375Max_T41=3453.906Throttle_Ratio=1.1Net_Thrust=53,008Sgross_ft=4535-------------------------------------FOM=3.43ROC=311 fpmFuselage Tank Vol=18,923 lbBal Field Distance=9905 ftW/S_takeoff=77.4 psf100% Vj=1891 fpsDe-rate Vj=1640 fps

Optimized Design (307)

Fan_PR=2.891406Overall_PR=36Extraction_Ratio=0.9984375Max_T41=3453.906Throttle_Ratio=1.1Net_Thrust=53,008Sgross_ft=4535-------------------------------------FOM=3.43ROC=311 fpmFuselage Tank Vol=18,923 lbBal Field Distance=9905 ftW/S_takeoff=77.4 psf100% Vj=1891 fpsDe-rate Vj=1640 fps

Max FOM Benchmark

• Both Optimization 1 and GTRR Compromise were resized to meet constraints • Both meet 55,000 ft “threshold” max cruise altitude limit

– Changed GTRR Compromise engine design cruise point to Mach 1.6, 55,000 ft (originally M1.8, 4

GTRR Compromise Engine

Fan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Vj=1375 fps

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Vj=1375 fps

Optimization improved on GTRR Compromise cycle at the expense of V

Optimization improved on GTRR Compromise cycle at the expense of V

Optimized Design (307)

Fan_PR=2.891406Overall_PR=36Extraction_Ratio=0.9984375Max_T41=3453.906Throttle_Ratio=1.1Net_Thrust=53,008Sgross_ft=4535-------------------------------------FOM=3.43ROC=311 fpmFuselage Tank Vol=18,923 lbBal Field Distance=9905 ftW/S_takeoff=77.4 psf100% Vj=1891 fpsDe-rate Vj=1640 fps

Optimized Design (307)

Fan_PR=2.891406Overall_PR=36Extraction_Ratio=0.9984375Max_T41=3453.906Throttle_Ratio=1.1Net_Thrust=53,008Sgross_ft=4535-------------------------------------FOM=3.43ROC=311 fpmFuselage Tank Vol=18,923 lbBal Field Distance=9905 ftW/S_takeoff=77.4 psf100% Vj=1891 fpsDe-rate Vj=1640 fps

Max FOM Benchmark
140 

Figure 3.4.2 765-072B Optimization #2 Summary (Minimize Vj) 

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Pwr=1375 fps

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Pwr=1375 fps

Optimization improved on GTRR Compromise cycle Vj and FOMCould not achieve 1100 fps Vj & other constraints with this NPSS engine cyc

Optimization improved on GTRR Compromise cycle Vj and FOMCould not achieve 1100 fps Vj & other constraints with this NPSS engine cyc

Min Vj Benchmark

Optimized Design (231)

Fan_PR=2.10723Overall_PR=34.4961Extraction_Ratio=1.05938Max_T41=3666.8Throttle_Ratio=1.12539Net_Thrust=64,688Sgross_ft=4949-------------------------------------FOM=3.17ROC=307 fpmFuselage Tank Vol=12,318 lbBal Field Distance=9970 ftW/S_takeoff=78.0 psf100% Vj=1397 fpsDe-rate Vj=1194 fps

Optimized Design (231)

Fan_PR=2.10723Overall_PR=34.4961Extraction_Ratio=1.05938Max_T41=3666.8Throttle_Ratio=1.12539Net_Thrust=64,688Sgross_ft=4949-------------------------------------FOM=3.17ROC=307 fpmFuselage Tank Vol=12,318 lbBal Field Distance=9970 ftW/S_takeoff=78.0 psf100% Vj=1397 fpsDe-rate Vj=1194 fps

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Pwr=1375 fps

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Pwr=1375 fps

Optimization improved on GTRR Compromise cycle Vj and FOMCould not achieve 1100 fps Vj & other constraints with this NPSS engine cyc

Optimization improved on GTRR Compromise cycle Vj and FOMCould not achieve 1100 fps Vj & other constraints with this NPSS engine cyc

Min Vj Benchmark

Optimized Design (231)

Fan_PR=2.10723Overall_PR=34.4961Extraction_Ratio=1.05938Max_T41=3666.8Throttle_Ratio=1.12539Net_Thrust=64,688Sgross_ft=4949-------------------------------------FOM=3.17ROC=307 fpmFuselage Tank Vol=12,318 lbBal Field Distance=9970 ftW/S_takeoff=78.0 psf100% Vj=1397 fpsDe-rate Vj=1194 fps

Optimized Design (231)

Fan_PR=2.10723Overall_PR=34.4961Extraction_Ratio=1.05938Max_T41=3666.8Throttle_Ratio=1.12539Net_Thrust=64,688Sgross_ft=4949-------------------------------------FOM=3.17ROC=307 fpmFuselage Tank Vol=12,318 lbBal Field Distance=9970 ftW/S_takeoff=78.0 psf100% Vj=1397 fpsDe-rate Vj=1194 fps
141 

Figure 3.4.3 765-072B Optimization #3 Summary (Maximize FOM/Vj) 

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Vj=1375 fps

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Vj=1375 fps

Optimization improved on GTRR Optimization improved on GTRR Compromise cycle Vj and FOM Compromise cycle Vj and FOM

Max FOM / Vj Benchmark

Optimized Design (244)

Fan_PR=2.0900039Overall_PR=36Extraction_Ratio=0.9738281Max_T41=3700Throttle_Ratio=1.109766Net_Thrust=65,508Sgross_ft=4930-------------------------------------FOM=3.22ROC= 312 fpmFuselage Tank Vol= 11,101 lbBal Field Distance= 9979 ftW/S_takeoff= 77.8 psf100% Vj= 1422 fpsDe-rate Vj=1204 fps

Optimized Design (244)

Fan_PR=2.0900039Overall_PR=36Extraction_Ratio=0.9738281Max_T41=3700Throttle_Ratio=1.109766Net_Thrust=65,508Sgross_ft=4930-------------------------------------FOM=3.22ROC= 312 fpmFuselage Tank Vol= 11,101 lbBal Field Distance= 9979 ftW/S_takeoff= 77.8 psf100% Vj= 1422 fpsDe-rate Vj=1204 fps

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Vj=1375 fps

GTRR Compromise EngineFan_PR=2.24754Overall_PR=35.0555Extraction_Ratio=0.91115Max_T41=3127.96Throttle_Ratio=1.14895Net_Thrust=61,000Sgross_ft=5000-------------------------------------FOM=3.14ROC=360 fpmFuselage Tank Vol=11,480 lbBal Field Distance=9868 ftW/S_takeoff=76.6 psf100% Vj=1600 fpsDe-rate Vj=1375 fps

Optimization improved on GTRR Optimization improved on GTRR Compromise cycle Vj and FOM Compromise cycle Vj and FOM

Max FOM / Vj Benchmark

Optimized Design (244)

Fan_PR=2.0900039Overall_PR=36Extraction_Ratio=0.9738281Max_T41=3700Throttle_Ratio=1.109766Net_Thrust=65,508Sgross_ft=4930-------------------------------------FOM=3.22ROC= 312 fpmFuselage Tank Vol= 11,101 lbBal Field Distance= 9979 ftW/S_takeoff= 77.8 psf100% Vj= 1422 fpsDe-rate Vj=1204 fps

Optimized Design (244)

Fan_PR=2.0900039Overall_PR=36Extraction_Ratio=0.9738281Max_T41=3700Throttle_Ratio=1.109766Net_Thrust=65,508Sgross_ft=4930-------------------------------------FOM=3.22ROC= 312 fpmFuselage Tank Vol= 11,101 lbBal Field Distance= 9979 ftW/S_takeoff= 77.8 psf100% Vj= 1422 fpsDe-rate Vj=1204 fps
142 

Figure 3.4.4 765-072B Optimization Comparison 

Max FOM had highest FOM value at 3.43Min Vj had lowest Vj value at 1194 fpsMax FOM / Vj was a compromise

Max FOM had highest FOM value at 3.43Min Vj had lowest Vj value at 1194 fpsMax FOM / Vj was a compromise

Min Vj Max FOM / V

Optimized Design (307)

Fan_PR=2.891406Overall_PR=36Extraction_Ratio=0.9984375Max_T41=3453.906Throttle_Ratio=1.1Net_Thrust=53,008Sgross_ft=4535-------------------------------------FOM=3.43ROC=311 fpmFuselage Tank Vol=18,923 lbBal Field Distance=9905 ftW/S_takeoff=77.4 psf100% Vj=1891 fpsDe-rate Vj=1640 fps

Optimized Design (307)

Fan_PR=2.891406Overall_PR=36Extraction_Ratio=0.9984375Max_T41=3453.906Throttle_Ratio=1.1Net_Thrust=53,008Sgross_ft=4535-------------------------------------FOM=3.43ROC=311 fpmFuselage Tank Vol=18,923 lbBal Field Distance=9905 ftW/S_takeoff=77.4 psf100% Vj=1891 fpsDe-rate Vj=1640 fps

Max FOM

Optimized Design (231)

Fan_PR=2.10723Overall_PR=34.4961Extraction_Ratio=1.05938Max_T41=3666.8Throttle_Ratio=1.12539Net_Thrust=64,688Sgross_ft=4949-------------------------------------FOM=3.17ROC=307 fpmFuselage Tank Vol=12,318 lbBal Field Distance=9970 ftW/S_takeoff=78.0 psf100% Vj=1397 fpsDe-rate Vj=1194 fps

Optimized Design (231)

Fan_PR=2.10723Overall_PR=34.4961Extraction_Ratio=1.05938Max_T41=3666.8Throttle_Ratio=1.12539Net_Thrust=64,688Sgross_ft=4949-------------------------------------FOM=3.17ROC=307 fpmFuselage Tank Vol=12,318 lbBal Field Distance=9970 ftW/S_takeoff=78.0 psf100% Vj=1397 fpsDe-rate Vj=1194 fps

Optimized Design (2

Fan_PR=2.0900039Overall_PR=36Extraction_Ratio=0.973Max_T41=3700Throttle_Ratio=1.10976Net_Thrust=65,508Sgross_ft=4930-------------------------------FOM=3.22ROC= 312 fpmFuselage Tank Vol= 11Bal Field Distance= 997W/S_takeoff= 77.8 psf100% Vj= 1422 fpsDe-rate Vj=1204 fps

Optimized Design (24

Fan_PR=2.0900039Overall_PR=36Extraction_Ratio=0.973Max_T41=3700Throttle_Ratio=1.10976Net_Thrust=65,508Sgross_ft=4930-------------------------------FOM=3.22ROC= 312 fpmFuselage Tank Vol= 11Bal Field Distance= 997W/S_takeoff= 77.8 psf100% Vj= 1422 fpsDe-rate Vj=1204 fps

Max FOM had highest FOM value at 3.43Min Vj had lowest Vj value at 1194 fpsMax FOM / Vj was a compromise

Max FOM had highest FOM value at 3.43Min Vj had lowest Vj value at 1194 fpsMax FOM / Vj was a compromise

Min Vj Max FOM / V

Optimized Design (307)

Fan_PR=2.891406Overall_PR=36Extraction_Ratio=0.9984375Max_T41=3453.906Throttle_Ratio=1.1Net_Thrust=53,008Sgross_ft=4535-------------------------------------FOM=3.43ROC=311 fpmFuselage Tank Vol=18,923 lbBal Field Distance=9905 ftW/S_takeoff=77.4 psf100% Vj=1891 fpsDe-rate Vj=1640 fps

Optimized Design (307)

Fan_PR=2.891406Overall_PR=36Extraction_Ratio=0.9984375Max_T41=3453.906Throttle_Ratio=1.1Net_Thrust=53,008Sgross_ft=4535-------------------------------------FOM=3.43ROC=311 fpmFuselage Tank Vol=18,923 lbBal Field Distance=9905 ftW/S_takeoff=77.4 psf100% Vj=1891 fpsDe-rate Vj=1640 fps

Max FOM

Optimized Design (231)

Fan_PR=2.10723Overall_PR=34.4961Extraction_Ratio=1.05938Max_T41=3666.8Throttle_Ratio=1.12539Net_Thrust=64,688Sgross_ft=4949-------------------------------------FOM=3.17ROC=307 fpmFuselage Tank Vol=12,318 lbBal Field Distance=9970 ftW/S_takeoff=78.0 psf100% Vj=1397 fpsDe-rate Vj=1194 fps

Optimized Design (231)

Fan_PR=2.10723Overall_PR=34.4961Extraction_Ratio=1.05938Max_T41=3666.8Throttle_Ratio=1.12539Net_Thrust=64,688Sgross_ft=4949-------------------------------------FOM=3.17ROC=307 fpmFuselage Tank Vol=12,318 lbBal Field Distance=9970 ftW/S_takeoff=78.0 psf100% Vj=1397 fpsDe-rate Vj=1194 fps

Optimized Design (2

Fan_PR=2.0900039Overall_PR=36Extraction_Ratio=0.973Max_T41=3700Throttle_Ratio=1.10976Net_Thrust=65,508Sgross_ft=4930-------------------------------FOM=3.22ROC= 312 fpmFuselage Tank Vol= 11Bal Field Distance= 997W/S_takeoff= 77.8 psf100% Vj= 1422 fpsDe-rate Vj=1204 fps

Optimized Design (24

Fan_PR=2.0900039Overall_PR=36Extraction_Ratio=0.973Max_T41=3700Throttle_Ratio=1.10976Net_Thrust=65,508Sgross_ft=4930-------------------------------FOM=3.22ROC= 312 fpmFuselage Tank Vol= 11Bal Field Distance= 997W/S_takeoff= 77.8 psf100% Vj= 1422 fpsDe-rate Vj=1204 fps
143 

Figure 3.5.1 765-076E Optimization #1 Summary (Maximize Range) 

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

• Optimization improved on Benchmark cycle for both range and Vj• Could not achieve threshold 4000 nm range

• Optimization improved on Benchmark cycle for both range and Vj• Could not achieve threshold 4000 nm range

Optimized Design (124)

Fan_PR=2.3541Overall_PR=38.0469Extraction_Ratio=0.97695Max_T41=4000Throttle_Ratio=1.0793Net_Thrust=41,016Range=3910-------------------------------------FOM=1.57TOGW=179,807 lbsROC=310 fpmExtra Fuel Tank Vol=3118 lbsBal Field Distance= 9905 ftW/S_takeoff=71.5 psfFan Diameter=73.4 in100% Vj= 1578 fpsDe-rate Vj=1250 fps

Optimized Design (124)

Fan_PR=2.3541Overall_PR=38.0469Extraction_Ratio=0.97695Max_T41=4000Throttle_Ratio=1.0793Net_Thrust=41,016Range=3910-------------------------------------FOM=1.57TOGW=179,807 lbsROC=310 fpmExtra Fuel Tank Vol=3118 lbsBal Field Distance= 9905 ftW/S_takeoff=71.5 psfFan Diameter=73.4 in100% Vj= 1578 fpsDe-rate Vj=1250 fps

Max Range Benchmark

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

• Optimization improved on Benchmark cycle for both range and Vj• Could not achieve threshold 4000 nm range

• Optimization improved on Benchmark cycle for both range and Vj• Could not achieve threshold 4000 nm range

Optimized Design (124)

Fan_PR=2.3541Overall_PR=38.0469Extraction_Ratio=0.97695Max_T41=4000Throttle_Ratio=1.0793Net_Thrust=41,016Range=3910-------------------------------------FOM=1.57TOGW=179,807 lbsROC=310 fpmExtra Fuel Tank Vol=3118 lbsBal Field Distance= 9905 ftW/S_takeoff=71.5 psfFan Diameter=73.4 in100% Vj= 1578 fpsDe-rate Vj=1250 fps

Optimized Design (124)

Fan_PR=2.3541Overall_PR=38.0469Extraction_Ratio=0.97695Max_T41=4000Throttle_Ratio=1.0793Net_Thrust=41,016Range=3910-------------------------------------FOM=1.57TOGW=179,807 lbsROC=310 fpmExtra Fuel Tank Vol=3118 lbsBal Field Distance= 9905 ftW/S_takeoff=71.5 psfFan Diameter=73.4 in100% Vj= 1578 fpsDe-rate Vj=1250 fps

Max Range Benchmark
144 

Figure 3.5.2 765-076E Optimization #2 Summary (Minimize Vj) 

Optimization improved on Benchmark cycle for Vj at the expense of rangeCould not achieve 1100 fps Vj & other constraints with this NPSS engine c

Optimization improved on Benchmark cycle for Vj at the expense of rangeCould not achieve 1100 fps Vj & other constraints with this NPSS engine c

Min Vj Benchmark

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Optimized Design (162)

Fan_PR=2.14609Overall_PR=20Extraction_Ratio=1.07773Max_T41=3906Throttle_Ratio=1.11875Net_Thrust=37,930Range=3500-------------------------------------FOM=1.51TOGW=175,598 lbsROC=316 fpmExtra Fuel Tank Vol=7597 lbsBal Field Distance= 9972 ftW/S_takeoff=69.8 psfFan Diameter=75.2 in100% Vj= 1437 fpsDe-rate Vj=1164 fps

Optimized Design (162)

Fan_PR=2.14609Overall_PR=20Extraction_Ratio=1.07773Max_T41=3906Throttle_Ratio=1.11875Net_Thrust=37,930Range=3500-------------------------------------FOM=1.51TOGW=175,598 lbsROC=316 fpmExtra Fuel Tank Vol=7597 lbsBal Field Distance= 9972 ftW/S_takeoff=69.8 psfFan Diameter=75.2 in100% Vj= 1437 fpsDe-rate Vj=1164 fps

Optimization improved on Benchmark cycle for Vj at the expense of rangeCould not achieve 1100 fps Vj & other constraints with this NPSS engine c

Optimization improved on Benchmark cycle for Vj at the expense of rangeCould not achieve 1100 fps Vj & other constraints with this NPSS engine c

Min Vj Benchmark

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Optimized Design (162)

Fan_PR=2.14609Overall_PR=20Extraction_Ratio=1.07773Max_T41=3906Throttle_Ratio=1.11875Net_Thrust=37,930Range=3500-------------------------------------FOM=1.51TOGW=175,598 lbsROC=316 fpmExtra Fuel Tank Vol=7597 lbsBal Field Distance= 9972 ftW/S_takeoff=69.8 psfFan Diameter=75.2 in100% Vj= 1437 fpsDe-rate Vj=1164 fps

Optimized Design (162)

Fan_PR=2.14609Overall_PR=20Extraction_Ratio=1.07773Max_T41=3906Throttle_Ratio=1.11875Net_Thrust=37,930Range=3500-------------------------------------FOM=1.51TOGW=175,598 lbsROC=316 fpmExtra Fuel Tank Vol=7597 lbsBal Field Distance= 9972 ftW/S_takeoff=69.8 psfFan Diameter=75.2 in100% Vj= 1437 fpsDe-rate Vj=1164 fps
145 

Figure 3.5.3 765-076E Optimization #3 Summary (Maximize Range/Vj) 

Optimization improved on Optimization improved on Benchmark cycle Vj and Range Benchmark cycle Vj and Range

Max Range / Vj Benchmark

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Optimized Design (423)

Fan_PR=2.20176Overall_PR=39.2188Extraction_Ratio=0.9Max_T41=4000Throttle_Ratio=1.10039Net_Thrust=41,133Range=3799-------------------------------------FOM=1.57TOGW=178,581 lbsROC=343 fpmExtra Fuel Tank Vol=4904 lbsBal Field Distance= 9986 ftW/S_takeoff=71.0 psfFan Diameter=75.3 in100% Vj= 1517 fpsDe-rate Vj=1197 fps

Optimized Design (423)

Fan_PR=2.20176Overall_PR=39.2188Extraction_Ratio=0.9Max_T41=4000Throttle_Ratio=1.10039Net_Thrust=41,133Range=3799-------------------------------------FOM=1.57TOGW=178,581 lbsROC=343 fpmExtra Fuel Tank Vol=4904 lbsBal Field Distance= 9986 ftW/S_takeoff=71.0 psfFan Diameter=75.3 in100% Vj= 1517 fpsDe-rate Vj=1197 fps

Optimization improved on Optimization improved on Benchmark cycle Vj and Range Benchmark cycle Vj and Range

Max Range / Vj Benchmark

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Starting Cycle ParametersFan_PR=2.89141Overall_PR=36Extraction_Ratio=0.99844Max_T41=3454Throttle_Ratio=1.1Net_Thrust=38,000Range=3700-------------------------------------FOM=1.51TOGW=179,872 lbsROC=345 fpmExtra Fuel Tank Vol= 4273 lbsBal Field Distance=9806 ftW/S_takeoff=71.5 psfFan Diameter=63.7 in100% Vj=1878 fpsDe-rate Vj=1510 fps

Optimized Design (423)

Fan_PR=2.20176Overall_PR=39.2188Extraction_Ratio=0.9Max_T41=4000Throttle_Ratio=1.10039Net_Thrust=41,133Range=3799-------------------------------------FOM=1.57TOGW=178,581 lbsROC=343 fpmExtra Fuel Tank Vol=4904 lbsBal Field Distance= 9986 ftW/S_takeoff=71.0 psfFan Diameter=75.3 in100% Vj= 1517 fpsDe-rate Vj=1197 fps

Optimized Design (423)

Fan_PR=2.20176Overall_PR=39.2188Extraction_Ratio=0.9Max_T41=4000Throttle_Ratio=1.10039Net_Thrust=41,133Range=3799-------------------------------------FOM=1.57TOGW=178,581 lbsROC=343 fpmExtra Fuel Tank Vol=4904 lbsBal Field Distance= 9986 ftW/S_takeoff=71.0 psfFan Diameter=75.3 in100% Vj= 1517 fpsDe-rate Vj=1197 fps
146 

Figure 3.5.4 765-076E Optimization Comparison 

Max Range had highest Range value at 3910 nmMin Vj had lowest Vj value at 1164 fpsMax Range / Vj was a compromise

Max Range had highest Range value at 3910 nmMin Vj had lowest Vj value at 1164 fpsMax Range / Vj was a compromise

Max Range Min Vj Max Range / V

Optimized Design (124)

Fan_PR=2.3541Overall_PR=38.0469Extraction_Ratio=0.97695Max_T41=4000Throttle_Ratio=1.0793Net_Thrust=41,016Range=3910-------------------------------------FOM=1.57TOGW=179,807 lbsROC=310 fpmExtra Fuel Tank Vol=3118 lbsBal Field Distance= 9905 ftW/S_takeoff=71.5 psfFan Diameter=73.4 in100% Vj= 1578 fpsDe-rate Vj=1250 fps

Optimized Design (124)

Fan_PR=2.3541Overall_PR=38.0469Extraction_Ratio=0.97695Max_T41=4000Throttle_Ratio=1.0793Net_Thrust=41,016Range=3910-------------------------------------FOM=1.57TOGW=179,807 lbsROC=310 fpmExtra Fuel Tank Vol=3118 lbsBal Field Distance= 9905 ftW/S_takeoff=71.5 psfFan Diameter=73.4 in100% Vj= 1578 fpsDe-rate Vj=1250 fps

Optimized Design (162)

Fan_PR=2.14609Overall_PR=20Extraction_Ratio=1.07773Max_T41=3906Throttle_Ratio=1.11875Net_Thrust=37,930Range=3500-------------------------------------FOM=1.51TOGW=175,598 lbsROC=316 fpmExtra Fuel Tank Vol=7597 lbsBal Field Distance= 9972 ftW/S_takeoff=69.8 psfFan Diameter=75.2 in100% Vj= 1437 fpsDe-rate Vj=1164 fps

Optimized Design (162)

Fan_PR=2.14609Overall_PR=20Extraction_Ratio=1.07773Max_T41=3906Throttle_Ratio=1.11875Net_Thrust=37,930Range=3500-------------------------------------FOM=1.51TOGW=175,598 lbsROC=316 fpmExtra Fuel Tank Vol=7597 lbsBal Field Distance= 9972 ftW/S_takeoff=69.8 psfFan Diameter=75.2 in100% Vj= 1437 fpsDe-rate Vj=1164 fps

Optimized Design (423

Fan_PR=2.20176Overall_PR=39.2188Extraction_Ratio=0.9Max_T41=4000Throttle_Ratio=1.10039Net_Thrust=41,133Range=3799----------------------------------FOM=1.57TOGW=178,581 lbsROC=343 fpmExtra Fuel Tank Vol=4904 Bal Field Distance= 9986 W/S_takeoff=71.0 psfFan Diameter=75.3 in100% Vj= 1517 fpsDe-rate Vj=1197 fps

Optimized Design (423)

Fan_PR=2.20176Overall_PR=39.2188Extraction_Ratio=0.9Max_T41=4000Throttle_Ratio=1.10039Net_Thrust=41,133Range=3799-----------------------------------FOM=1.57TOGW=178,581 lbsROC=343 fpmExtra Fuel Tank Vol=4904 Bal Field Distance= 9986 fW/S_takeoff=71.0 psfFan Diameter=75.3 in100% Vj= 1517 fpsDe-rate Vj=1197 fps

Max Range had highest Range value at 3910 nmMin Vj had lowest Vj value at 1164 fpsMax Range / Vj was a compromise

Max Range had highest Range value at 3910 nmMin Vj had lowest Vj value at 1164 fpsMax Range / Vj was a compromise

Max Range Min Vj Max Range / V

Optimized Design (124)

Fan_PR=2.3541Overall_PR=38.0469Extraction_Ratio=0.97695Max_T41=4000Throttle_Ratio=1.0793Net_Thrust=41,016Range=3910-------------------------------------FOM=1.57TOGW=179,807 lbsROC=310 fpmExtra Fuel Tank Vol=3118 lbsBal Field Distance= 9905 ftW/S_takeoff=71.5 psfFan Diameter=73.4 in100% Vj= 1578 fpsDe-rate Vj=1250 fps

Optimized Design (124)

Fan_PR=2.3541Overall_PR=38.0469Extraction_Ratio=0.97695Max_T41=4000Throttle_Ratio=1.0793Net_Thrust=41,016Range=3910-------------------------------------FOM=1.57TOGW=179,807 lbsROC=310 fpmExtra Fuel Tank Vol=3118 lbsBal Field Distance= 9905 ftW/S_takeoff=71.5 psfFan Diameter=73.4 in100% Vj= 1578 fpsDe-rate Vj=1250 fps

Optimized Design (162)

Fan_PR=2.14609Overall_PR=20Extraction_Ratio=1.07773Max_T41=3906Throttle_Ratio=1.11875Net_Thrust=37,930Range=3500-------------------------------------FOM=1.51TOGW=175,598 lbsROC=316 fpmExtra Fuel Tank Vol=7597 lbsBal Field Distance= 9972 ftW/S_takeoff=69.8 psfFan Diameter=75.2 in100% Vj= 1437 fpsDe-rate Vj=1164 fps

Optimized Design (162)

Fan_PR=2.14609Overall_PR=20Extraction_Ratio=1.07773Max_T41=3906Throttle_Ratio=1.11875Net_Thrust=37,930Range=3500-------------------------------------FOM=1.51TOGW=175,598 lbsROC=316 fpmExtra Fuel Tank Vol=7597 lbsBal Field Distance= 9972 ftW/S_takeoff=69.8 psfFan Diameter=75.2 in100% Vj= 1437 fpsDe-rate Vj=1164 fps

Optimized Design (423

Fan_PR=2.20176Overall_PR=39.2188Extraction_Ratio=0.9Max_T41=4000Throttle_Ratio=1.10039Net_Thrust=41,133Range=3799----------------------------------FOM=1.57TOGW=178,581 lbsROC=343 fpmExtra Fuel Tank Vol=4904 Bal Field Distance= 9986 W/S_takeoff=71.0 psfFan Diameter=75.3 in100% Vj= 1517 fpsDe-rate Vj=1197 fps

Optimized Design (423)

Fan_PR=2.20176Overall_PR=39.2188Extraction_Ratio=0.9Max_T41=4000Throttle_Ratio=1.10039Net_Thrust=41,133Range=3799-----------------------------------FOM=1.57TOGW=178,581 lbsROC=343 fpmExtra Fuel Tank Vol=4904 Bal Field Distance= 9986 fW/S_takeoff=71.0 psfFan Diameter=75.3 in100% Vj= 1517 fpsDe-rate Vj=1197 fps
147 

Figure 3.5.5 CCA Diagram 
148 

4.0 Non-Proprietary Small Supersonic Aircraft 

4.1 Concept Selection 

The 076E was selected as the deliverable small supersonic aircraft because of its low boom characteristics. 

The mission performance rules are contained in Figure 4.1.1. The performance of the concept is shown in 

Figure 4.1.2 

4.2 Description of the 076E 

The model 765-076E was conceived as compromise configuration between sonic boom mitigation and 

airplane productivity. Some of the high level features of the configuration are shown in Figure 4.2.1. Since 

one of the primary drivers of sonic boom noise level is weight, the passenger capacity of the airplane was 

limited to 15-45. The performance of the airplane was analyzed with 30 passengers. 

Configuration Description 

The general arrangement of the aircraft is shown in Figure 4.2.2. An IGES file for this outer mold line is 

delivered with this report in a separate file titled “765-076E_ContractDeliverable.igs. A CFD quality OML 

IGES file is also included and it is titled “t65-076E_CFD_Ready.igs. 

The model 765-076E has a highly integrated arrangement with heavily shaped surfaces for improved 

supersonic drag and noise characteristics. The surface definition was created from a required area 

distribution to achieve a desired sonic boom, with sufficient constraints to ensure the ability to package 

internal components after the fact. The linear analysis that created the lofted geometry produced surfaces 

sufficient for the level of analysis done under this contract, but with design refinement the predicted sonic 

boom and cruise performance goals should be achievable with higher order analysis tools and experimental 

validation. 

Wing 

The wing has a highly swept planform with a large glove and aft deck to accommodate the integration of 

the upper mounted engine nacelles. The wing planform is shown in Figure 4.2.3. 

The outboard wing is of solid titanium construction. The inboard wing is of built-up titanium construction. 

An overview of the wing structure can be seen in the general arrangement drawing. The wing uses a 

drooped leading edge, midboard flaps, and an aileron, as well as two aft elevator surfaces. 

Tail 

The tail uses similar construction as the wing, with the outboard portion constructed of solid titanium, and 

the inboard portion constructed of built-up titanium. An overview of the tail structure can be seen in the 

general arrangement drawing. The tail is all-flying and is controlled by an actuator attached to the rear spar, 

turning about a spindle attached to the front spar. 

Body 

The body does not have a constant section and contains the passengers and crew, as well as baggage, 

landing gear, and subsystems. The body is constructed of built up titanium. The forward pressure bulkhead 

is slanted to allow the nose landing gear to be mounted further aft. Frames are distributed through the body 

on nominal 20 inch spacing on center. There are 28, 10 inch diameter passenger windows. A conceptual 

structural arrangement can be seen on the general arrangement drawing. 
149 

Landing Gear 

The landing gear is a conventional tricycle type. The main landing gear has 2 struts, with 6 wheels per strut. 

The truck is levered (similar to 777) such that the airplane rotates about the aft tire. The tires are high 

pressure (320 psi) and are 21.5 x 9. The MLG folds inboard and forward into the wing / body fairing. Two 

diagonal braces replace the traditional drag and sway braces to accommodate the shorter trunion. The nose 

landing gear has 2 tires which are 20 x 9. The nose landing gear folds forward in front of the forward 

pressure bulkhead, so there is no NLG pressure deck. 

Systems 

The flight deck is a conventional side-by-side arrangement with Pilot and Copilot. No provisions are made 

for a drooping nose or forward windscreen, so synthetic vision will be required. Windows to the side are 

provided, so synthetic vision will only be required to the front. The cockpit vision provided is shown in 

Figure 4.2.4. 

The electronics and avionics are located in the E/E bay behind the aft pressure bulkhead. Behind that are 

the ECS packs, and behind that the APU. A layout of the systems bays can be seen in Figure 4.2.5. 

The fuel is distributed among 10 tanks throughout the inboard wing and aft fuselage. A layout of the fuel 

tanks can be seen in Figure 4.2.6. Fuel can pumped from tank-to-tank to actively manage the airplane 

center of gravity for improved performance and reduced sonic boom. The fuel tanks can accommodate a 

total of 85,300 lb of fuel. 

Propulsion 

The propulsion system was developed as a part of contract Option 2, and will be discussed in that portion 

of the report. 

Payloads 

As many as 45 passengers can be accommodated on the 765-076E. Two notional Layouts of Passenger 

Accommodations (LOPAs) can be seen in Figure 4.2.7. Four type 1 emergency exits are provided, 2 in the 

front cabin, and 2 over the wing. The aft door cannot be serviced without walking on the wing, so the cabin 

will be serviced from the forward door only. 

Weights 

The 765-076E aircraft was sized to 180,000 lbs Maximum Takeoff Gross Weight (MTOGW) in order to 

meet low-boom and performance targets. The Operational Empty Weight (OEW) was calculated using the 

QWIKO weight estimation tool; a parametric supersonic aircraft weight estimation tool that was developed 

jointly with Boeing and McDonnell Douglas during the HSR program. QWIKO has been used successfully 

on several supersonic aircraft studies in the past, including the High Speed Civil Transport and Quiet 

Supersonic Platform studies. QWIKO includes allowances for N+2 technology and weight savings. 

The initial QWIKO weight statement is shown in Figure 4.2.8. This weight statement utilizes a BSE 

engine. The fuel weight was based on the calculated OEW and the design TOGW. QWIKO assumes that 

the structure is laid out efficiently and that there is sufficient structural depth to carry loads. The 765-076E 

configuration is very thin structurally in order to meet the low-boom requirements and preliminary FEM 

results indicate that the 180,000lb TOGW target may be unrealistic and very difficult to reach. This is 

discussed further in the Option 2 FEM section. 

A preliminary center of gravity diagram was constructed for the 765-076E configuration as shown in 

Figure 4.2.9. Due to the integrated nature of a low-boom configuration, the 765-076E was analyzed “as-is” 

and there was no additional tail sizing or wing shifting performed. 
150 

The Operational Empty Weight center of gravity was determined using QWIKO and by geographically 

locating large components such as the APU and EE Bay. The nose landing gear was located forward of the 

forward slant bulkhead and the main landing gear was placed as far aft as possible without interfering with 

the wings rear spar. This gear layout resulted in approximately 2% of the OEW on the nose wheel, and 5% 

of MTOGW on the nose wheel. 

During supersonic flight the center of pressure moves aft, so separate C.G. envelopes were needed for 

supersonic and subsonic flight. The aft limit was notionally set at 2% Mean Aerodynamic Chord (MAC) 

ahead of the main landing gear. The subsonic forward limit was set by the Maximum Zero Fuel Weight 

(MZFW) of the forward-most payload case, 30 passengers. A notional supersonic forward limit was set 5% 

in front of the aft limit, which is comparable to the High Speed Civil Transport’s envelope. 

The preliminary 765-076E weights were revised to include the findings from the NPSS engine study and 

initial performance runs. The TOGW was reduced to 178,737 lbs and the OEW was recalculated using 

QWIKO. The 765-076E performance weight statement can be found in Figure 

4.2.10. 

4.3 Sonic Boom Discussion 

The 765-076E Signature was designed using the Boeing MTA tool. It was based on a linear design and the 

signature was analyzed with MDBOOM and Zephyrus wave propagation codes. Zephyrus results are 

believed to be more accurate, because it applies molecular relaxation throughout the wave propagation. 

The MDBOOM code is utilized in off-body signature analysis, because it is easier to use. The wave 

propagation results are shown in Figure 4.3.1. The Zephyrus code yields a ground signature of 91.5 PLdB, 

while the MDBOOM code yields a ground signature of 93.8 PLdB. The 765-076E configuration should be 

able to provide ~91 PLdB after Non-linear CFD-based boom optimization. 

The off-body Euler CFD was conducted to see how effective the linear design was in shaping the front and 

aft signature. Figure 4.3.2 shows the near-field pressure distribution produced by the Cart3d CFD code. A 

comparison of the off-body F-Function, Cart3d vs. the linear theory, is shown in Figure 4.3.3. Note that the 

Cart3d & linear theory agree on some prominent features. Cart3d captures the wave advance neglected by 

Linear Theory. It spreads the length and steepens the shocks. The ground signature is shown in Figure 

4.3.4 for the original design condition, Mach=1.8, weight=162,000 lbs, and altitude=49,000 ft. Other 

conditions analyzed, included Mach=1.6 & 1.8 at altitudes=49,000 and 55,000 ft. The ground signature 

results for these alternate conditions can be found in Figures 4.3.5, 4.3.6, and 4.3.7. Examining the figures 

shows that there is some signature shaping on the front side of the sonic boom signature, but essentially an 

N-wave feature for the aft side of the sonic boom signature. The progression of one of the signatures with 

altitude is shown in Figure 4.3.8 to further illustrate the off-body CFD results. These signature differences 

(CFD vs. linear) prevents using only the linear tools for design. So the process is to develop concepts and 

seeds in linear theory and then optimize with Cart3d to realize the linear theory level. 

In summary the 765-076E was designed using linear-based methods (Boeing MTA) to provide ~91 PLdB 

shaped sonic-boom signature. Off-body CFD-based signature analysis (Cart3d Euler) shows a partially 

shaped front signature and an N-wave aft signature with ~ 97 – 100 PLdB. These results are typical and 

expected for an initial design without non-linear CFD-based boom optimization. Non-linear CFD-based 

boom optimization will be conducted in a separate N+2 System Level Experimental Validation contract. 

4.4 Goal Compliance 

The N+2 study identified some guidelines or goals for a variety of metrics. These are summarized in Table 

4.4.1. The capabilities of the two study concepts (-072B and -076E) are also shown in the table. This table 

shows that the -072B will meet the performance goals, but will not meet the environmental goals. While 
151 

the -076E has a lower boom signature, it still does not meet the study’s environmental goals. It also does 

not meet the minimum range or FOM goals of the study. 

4.5 Propulsion 

The propulsion system used for the mission analysis was a NPSS optimized cycle as described in section 5. 
152 

Tables and Figures for Section 4.0 

Standard N+2 Mission

Figure 4.1.1 765-076E Mission Profile

z Nominal Performancez Standard Dayz Fuel Density: 6.7 lb/US Gallon

Mission

Still Air Range

Block Time & Fuel

Climb

and 

Accelerate 

to 

Optimum

Supersonic

Cruise

M

&

h

Climbout and Accelerate

to 1,500 ft & 250 kts

Taxi Out (9 Minutes)

Takeoff to 35 ft

Approach and Land (4 Minutes)

Taxi In (5 Minutes from Reserves)

SupersonicCruise / ClimbConstant Mach, Optimum Altitude

(55k ft limit)

See specified 

climb / descent 

Schedule to right

Descend

and

decelerate

200 nm

Economy Climb

Cruise at LRC Mach

Approach & Land (4 Minutes)

30 Minute Hold @ 1,500 ft

Reserves

Zero Wind

5% Block Fuel Allowance

Missed Approach

Descend to 1,500 ft

z Nominal Performancez Standard Dayz Fuel Density: 6.7 lb/US Gallon

Mission

Still Air Range

Block Time & Fuel

Climb

and 

Accelerate 

to 

Optimum

Supersonic

Cruise

M

&

h

Climbout and Accelerate

to 1,500 ft & 250 kts

Taxi Out (9 Minutes)

Takeoff to 35 ft

Approach and Land (4 Minutes)

Taxi In (5 Minutes from Reserves)

SupersonicCruise / ClimbConstant Mach, Optimum Altitude

(55k ft limit)

See specified 

climb / descent 

Schedule to right

Descend

and

decelerate

200 nm

Economy Climb

Cruise at LRC Mach

Approach & Land (4 Minutes)

30 Minute Hold @ 1,500 ft

Reserves

Zero Wind

5% Block Fuel Allowance

Missed Approach

Descend to 1,500 ft
153 

Figure 4.1.2 765-076E Non-Proprietary Small Supersonic Aircraft Deliverable 

Performance Summary 

ConfigurationWing Area (ESDU)Wing Area (Gross)Thrust per engineFan Diameter (in)Engine WeightTOGWOEWPayload Total Fuel Block FuelRangeSupersonic Cruise

MachAltitudeWeightCLL/DSFC

Climb (pinch point)

MachAltitudeWeightCLL/DThrustSFCROC

De-rate Vj (10k ft takeoff)W/S (takeoff)FOM

Optimization 3

Maximum Range / Vj

NPSS GTRR Cycle

25173344

41,133

75.3

7271

178,581

88,390

6,300

84,37372,740

3799

1.6

51,784

161,165

0.1726

8.72

0.985

1.13

40,550

170.066

0.1985

8.551

20,7390.8096

343

1197

71.01.57

Preferred

Mach 1.630 PAXvariable range

ConfigurationWing Area (ESDU)Wing Area (Gross)Thrust per engineFan Diameter (in)Engine WeightTOGWOEWPayload Total Fuel Block FuelRangeSupersonic Cruise

MachAltitudeWeightCLL/DSFC

Climb (pinch point)

MachAltitudeWeightCLL/DThrustSFCROC

De-rate Vj (10k ft takeoff)W/S (takeoff)FOM

Optimization 3

Maximum Range / Vj

NPSS GTRR Cycle

25173344

41,133

75.3

7271

178,581

88,390

6,300

84,37372,740

3799

1.6

51,784

161,165

0.1726

8.72

0.985

1.13

40,550

170.066

0.1985

8.551

20,7390.8096

343

1197

71.01.57

Preferred

Mach 1.630 PAXvariable range
154 

Figure 4.2.1 Model 765-076E Features
155 

Figure 4.2.2 765-076E General Arrangement, 1/200 scale. 
156 

Figure 4.2.3 076E Wing Planform 

556.0790

0.0000

1278.6580

234.7800

1622.0920

503.1000

1636.9920

516.5160

1522.1010193.6940

SOB LE 50.0000

1159.4152

149.5693
157 

Figure 4.2.4 076E Cockpit Vision 

-50-25

0

2550

-150 -125 -100 -75 -50 -25 0 25 50 75 100 

Azimuth (deg)

Elevation (deg)

Available 

vision

Required 

vision
158 

Figure 4.2.5 076E Subsystem Layout 

Aft Pressure Bulkhead

E/E Bay

Environmental Control System

Auxiliary Power Unit
159 

Figure 4.2.6 076E Fuel Tank Layout 
160 

Figure 4.2.7 076E Layout of Passenger Accommodations 

422

Carts

0.1330.0830.333

Cart Ratio

-

24

6

Lav Ratio

Total - - 

Economy 80 1 

First 20 7 

Class % Closet Rati

422

Carts

0.1330.0830.333

Cart Ratio

-

24

6

Lav Ratio

Total - - 

Economy 80 1 

First 20 7 

Class % Closet Rati

321

Carts

0.0730.0510.250

Cart Ratio

-

39

4

Lav Ratio

Total - - 

Economy 90.2 0 

First 9.8 2 

Class % Closet Rati

321

Carts

0.0730.0510.250

Cart Ratio

-

39

4

Lav Ratio

Total - - 

Economy 90.2 0 

First 9.8 2 

Class % Closet Rati
161 

Figure 4.2.8 Preliminary 765-076E Performance Weights 

Growth 180,000 lb

FC Description Weight CG

01 Wing Structure 26670 1409

02 Horizontal Tail Struture 3110 1667

03 Vertical Tail Structure 10 1667

04 Fuselage Structure 9600 911

05 Main Landing Gear 4330 1285

07 Nose Landing Gear 620 241

08 Forebody Controls 0 0

Structure Total 44,340 1,291

09 Inlet Structure and Systems 2840 1458

10 Cowling 1560 1590

11 Pylon/Strut 1040 1583

12 Engine 19220 1567

13 Nozzle 1020 1715

14 Installation (incl. fairings) 1360 1583

15 Engine Accessories, Controls, & Start System 200 1605

Propulsion Pod Total 27,240 1,564

23 Fuel System 2500 1306

24 APU/EPU 450 1577

24 Instruments 830 459

24 Surface Controls 2010 1240

24 Hydraulic Power System 1300 1187

24 Pneumatic System 0 0

24 Electrical System 1920 1474

24 Electronics 510 1294

24 Flight Provisions 760 320

24 Passenger Accomodations 6010 841

24 Cargo Compartment 550 1368

24 Emergency Equipment 180 818

24 Environmental Control Systems 980 1262

24 Ice Protection 200 1424

24 Exterior Markings 380 1168

24 Load and Handling 0 0

24 Customer Options 800 500

Systems & Fixed Equipment Total 19,380 1,055

Manufacturer's Empty Weight (MEW) 90,960 1,323

97 Standard and Operational Items 3140 891

Operational Empty Weight (OEW) 94,100 1,308

Payload 6300 879

Max Zero Fuel Weight (MZFW) 100,400 1281

Fuel 79,600 1328

MTOGW 180,000 1,302
162 

Figure 4.2.9 Preliminary 765-076E Center of Gravity Envelope

765-076E Center of Gravity Envelope

85,00095,000

105,000115,000125,000135,000145,000155,000165,000175,000185,000

-20,000,000 -10,000,000 0 10,000,000 

Moment about 42% MAC (Sta 1288.78), in-lb

Aircraft Weight, lb

Main Landing Gear

(Middle wheel)

47.8% MAC

Sta 1333

2.5 G 

MZFW (94,720)

Fwd Lim - ZFW

30 Pax

32% MAC <--- 2% MAC 2% MAC ----> 42% MAC 50% MAC

Aft Lim - ZFW

Aux Ballast Tank

M=1

M=1.6

Start Cruise

M=1.6

End Cruise

Ferry Mission

60% Payload

60% Payload AFTFWD

Subsonic 

FWD LIMIT

Supersonic 

Envelope

MAC: STA 759

LEMAC: STA 970

OWE: 88,420 lb @ STA 1292

Fwd Limit: 41 %MAC

or STA 1281.2

Aft Limit: 46 %MAC

or STA 1319.1
163 

Figure 4.2.10 765-076E Performance Weights 

FC Description Weight CG

01 Wing Structure 26590 1409

02 Horizontal Tail Struture 3110 1668

03 Vertical Tail Structure 10 0

04 Fuselage Structure 9590 911

05 Main Landing Gear 4290 1285

07 Nose Landing Gear 620 241

08 Forebody Controls 0 0

Structure Total 44,210 1,291

09 Inlet Structure and Systems 2840 1458

10 Cowling 1560 1590

11 Pylon/Strut 760 1582

12 Engine 14540 1567

13 Nozzle 960 1715

14 Installation (incl. fairings) 1040 1582

15 Engine Accessories, Controls, & Start System 200 1605

Propulsion Pod Total 21,900 1,563

23 Fuel System 2420 1306

24 APU/EPU 450 1577

24 Instruments 830 459

24 Surface Controls 2000 1240

24 Hydraulic Power System 1290 1187

24 Pneumatic System 0 0

24 Electrical System 1810 1476

24 Electronics 510 1294

24 Flight Provisions 760 320

24 Passenger Accomodations 6010 841

24 Cargo Compartment 550 1368

24 Emergency Equipment 180 818

24 Environmental Control Systems 980 1262

24 Ice Protection 200 1424

24 Exterior Markings 380 1168

24 Load and Handling 0 0

24 Customer Options 800 500

Systems & Fixed Equipment Total 19,170 1,052

Manufacturer's Empty Weight (MEW) 85,280 1,307

97 Standard and Operational Items 3140 891

Operational Empty Weight (OEW) 88,420 1,292

Payload 6300 879

Max Zero Fuel Weight (MZFW) 94,720 1265

Fuel 84,017 1311

MTOGW 178,737 1,286
164 

Figure 4.3.1 765-076E As-Designed Shaped Ground Signatures 
165 

Figure 4.3.2 765-076E CART3D CFD Euler Solution 

Mach = 1.8 h = 49,000 ft. Wt. = 162,000 lbs. CL = 0.1117 

PLdB = 100.0 
166 

Figure 4.3.3 765-076E Off-Body F-Function CART3D vs. Linear
167 

Figure 4.3.4 765-076E Ground Signatures

Figure 4.3.5 765-076E Alternate Ground Signature Results 

 

Mach=1.8, Altitude=49 Kft, CL=0.1111, Weight=162,000 lbs

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

0.00 0.05 0.10 0.15 0.20 0.25 0.30 0.35 0.40 0.45

Time(sec)

Overpressure, 

ΔP (psf)

R/L = 2.0, PLdB = 100.5

R/L = 2.5, PLdB = 99.9

R/L = 3.0, PLdB = 100.0

Mach=1.6, Altitude=49 Kft, CL=0.1406, Weight=162,000 lbs

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

0.00 0.05 0.10 0.15 0.20 0.25 0.30 0.35 0.40 0.45 0.50

Time(sec)

Overpressure, 

ΔP (psf)

R/L = 2.0, PLdB = 99.7

R/L = 2.5, PLdB = 99.3

R/L = 3.0, PLdB = 99.9
168 

Figure 4.3.6 765-076E Alternate Ground Signature Results 

Figure 4.3.7 765-076E Alternate Ground Signature Results

Mach=1.8, Altitude=55 Kft, CL=0.1480, Weight=162,000 lbs

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

0.00 0.05 0.10 0.15 0.20 0.25 0.30 0.35 0.40 0.45 0.50

Time(sec)

Overpressure, 

ΔP (psf)

R/L = 2.0, PLdB = 100.3

R/L = 2.5, PLdB = 100.1

R/L = 3.0, PLdB = 100.5

Mach=1.6, Altitude=55 Kft, CL=0.1873, Weight=162,000 lbs

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

0.00 0.10 0.20 0.30 0.40 0.50 0.60

Time(sec)

Overpressure, 

ΔP (psf)

R/L = 2.0, PLdB = 97.8

R/L = 2.5, PLdB = 97.6

R/L = 3.0, PLdB = 98.3
169 

Figure 4.3.8 Progression of Signature with Altitude(Mach=1.6, Alt=49 kft, Weight=162 klbs) 

 

-5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf)

-5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf)

R/L = 3

Alt=48Kft

Alt=47Kft

Alt=46Kft

Alt=40Kft

Alt=30Kft

Alt=20Kft

Alt=10Kft

Ground

(1.9 Ref. Factor)

-5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf)

-5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf) -5.0

0.0

5.0

0 0.05 0.1 0.15 0.2

Time (sec)

ΔP (psf)

R/L = 3

Alt=48Kft

Alt=47Kft

Alt=46Kft

Alt=40Kft

Alt=30Kft

Alt=20Kft

Alt=10Kft

Ground

(1.9 Ref. Factor)
170 

Table 4.4.1 N+2 Study Compliance Matrix 

Guidelines for Study 765-072B 765-076E 

Environmental Goals 

Sonic Boom 

(Linear Theory Goal) 

65-70 PLdB ~ 100 ~ 90 

Airport Noise 

(cum below Stage 3) 

Stage 3 –10 to 20 dB 

(1100 fps Vj) 

1194 Vj minus 15 

Cruise Emissions 

(NOx g/kg of fuel) 

Limited Consideration Not evaluated Not evaluated 

Performance Goals 

Cruise Mach 1.6 to 2.0 1.6 1.6 

Range (nm) 4000 to 6000 4000 3799 

Payload 

(passengers) 

25 to 100 100 30 

Fuel Efficiency 

(passenger-miles per 

pound of fuel) 

>3.0 3.17 1.57 

 
171 

5.0 Second Iteration Analysis of Baseline Concept 

5.1 Structural Analysis (Option 1) 

The structural analysis and sizing of the selected configuration begins with the OML geometry and basic 

structural layout provided by the configuration design engineer. A finite element model (FEM) was 

developed based on this data utilizing Boeing ‘rapid FEM’ development tools. Initial sizing was estimated 

and material properties obtained for the baseline materials of carbon composites and aluminum selected for 

this analysis. A 2nd cycle of analysis was executed to improve the loads paths, evaluate alternate material 

usage and revised criteria. 

The flight envelope is shown in Figures 5.1.1. The dive speed envelope was defined as 1.15% of the cruise 

speed envelope. A limited set of conceptual design level structural design criteria was developed for initial 

sizing of the airframe primary components. These consisted of symmetrical pull-up conditions, steady roll, 

3–point landing (@ 1.6 g), 0.5 g ground turn, and cabin pressurization. The cabin is assumed to be 

pressurized to a design pressure altitude of 6000 Ft mean sea level. It is assumed that the pressure relief 

valve is set to 0.5 psi. Figure 5.1.2 lists the flight load conditions. Loads were developed for the target 

MTOW condition of 180,000 lbs. 

A weights statement together with parametric weight was prepared by the weight and balance group and 

shown in section 4.0. The weights group provided the distribution of the nonstructural weight items to the 

FEM for use in the loads and structural analyses 

A set of design loads conditions were generated from the initial FEM and used in the sizing analysis using 

MSCNASTRAN Solution 200. The final sizing analysis has produced structural weights that exceed the 

parametric weight estimates. The Maximum Take-Off Weight (MTOW) has been defined as 180,000 lbs 

and therefore the maximum fuel load has been reduced commensurate with the difference in the parametric 

structural weight estimates and the sized structure. The reduction in fuel load results in a reduction in 

mission range that falls below minimum requirements and decreases the value of the final sized FEM for 

use in further analyses. 

The following subsections provide additional detail and results for the structural analysis activities. 

5.1.1 FEM and Sizing 

Challenges 

Due to the unique configuration of the 765-076E airplane, there are a number of structural challenges for 

the team to address. Two of these challenges are the thin aft deck and the outboard and inboard wing joint. 

• Thin Aft deck, Figure 5.1.3: The aft deck configuration presents a significant challenge for the 

structure team, as it averages only 7 inches thick and require to carry the engine and the tail load. 

• Outboard and Inboard Wing Joint, Figure 5.1.4: The joint between the outboard and inboard wing 

also presented a significant structural challenge, as the depth is about 7 inches in that area. 

Assumptions 

The structural analysis task under this NRA is a low-fidelity analysis on a conceptual design model, 

therefore a number of assumptions had to be made for the analysis. The following is a list of assumptions 

used in this analysis: 

• The tails will act as all moving control surfaces. This assumption is made because the control 

surfaces and the control laws have not been fully developed, and the fact that a huge tail load is 

required to perform the pitch maneuver 
172 

• 10% improvement for material stress and strain allowable. This assumption is made to reduce 

vehicle weight and is deemed valid because the vehicle is designed to be a 2025 airplane 

• Assuming the control surface actuators will fit into the limited space in the airplane 

• The sizing analysis is a strength only analysis, so no buckling or dynamic constraints were 

considered 

• Control surfaces will not be sized. They are model to transfer the aero loads to the structure, and 

has weight assigned by the parametric tool 

Materials Used in the Finite Element Model 

• Sandwich Panel: 

E11 

(ksi) 

E22 

(ksi) 

Poisson 

Ratio 

Shear Modulus 

(ksi) 

Density 

(lb/in^3) 

Carbon Fiber Face 

Sheet 23000 1600 0.34 800 0.056 

Aluminum Honeycomb 187 0.187 0.4 0.013 0.00329 

• Isotropic Materials: 

E (ksi) Poisson Ratio 

Density 

(lb/in^3) 

Titanium 10000 0.33 0.101 

Aluminum 16000 0.31 0.162 

Carbon-Carbon 18460 0.34 0.065 

• The application of materials can be seen in the Figures 5.1.5 & 5.1.6: 
173 

Finite Element Model 

Structural Layout 

The structural layout of the 076E airplane is shown in Figures 5.1.7. This layout is the result of previous 

sizing analysis cycle, where the high running loads indicated issues with the aft deck stiffness and outboard 

and inboard wing joint. This can be seen in plots of the running loads, as shown in Figures 5.1.8 & 5.1.9. 

To enhance the structure, a number of improvements were added to the vehicle: 

• Tail torque box: Because of thin aft deck, a torque box was used to carry the tail load forward to 

the trailing edge of the wing. This is to relieve the load in the carry through structure in the aft 

deck, as the depth there are insufficient to carry the tail load and the engine weight effectively. 

• Keel structure, Figure 5.1.10: A set of keels were added to the inboard wing in the streamwise 

direction to add bending stiffness to the wing. This addition was deem necessary after initial sizing 

analysis indicating the need to increase the rigidity of the inboard wing and aft deck 

• Aft deck elevator surfaces: A set of elevator surfaces were added to the aft deck to relieve the tail 

load 

• Solid wing and tail tips: Solid core were used at the tip of the outboard wing and tail. This is due 

to the limited internal volumes at these locations. 

Structural Analysis Process 

The structural analysis process, shown schematically in Figure 5.1.11, utilized a variety of COTS tools as 

well as Boeing developed processes. The analysis began with an OML geometry of the airplane, which was 

then taken into Catia V5 to prepare for finite element modeling. The preparation includes the creation of 

internal structural layout, and the payload structure in the fuselage. All the surfaces are also split to make 

the meshing process easier. Once that is completed, it was taken into Patran to create finite element and 

aeroelastic model. A number of Boeing process was used to rapidly prepare the FEM for the loads and 

structural sizing analysis 

Setup for Sizing Analysis 

The sizing of the 076E FEM utilized SOL200 solver from MSC Nastran 2005R3. The objective of the 

sizing analysis is to find global minimum for weight based on strength sizing. There are a total of 1096 

design variables in the model, which is high for a conceptual design model. This is needed to try to achieve 

the aggressive weight target, which was calculated by Boeing’s parametric weight tool to be 31,430 

pounds. These design variables include the following: 

• Ply thickness of composite panels 

• Shell thickness of titanium structures 

• Cross sectional area of caps 

The sizing analysis was based on the strength of the vehicle, so strain allowables were used as the 

constraints for the panels and stress allowables were used as constraints for the caps. The sizing analysis 

performed here was based strictly on the linear static analyses of the vehicle, so other considerations such 

as flutter, buckling, and dynamic gust will need to be studied to yield the final weight and structure design. 

The load for each load cases (see table below) were first generated in Nastran SOL144 static aeroelastic 

analysis, and then mapped to Nastran FORCE cards. These FORCE cards are then combined with other 

load conditions if necessary, such as inertial load for the landing and the pressure loads for the fuselage. 

The usages of FORCE cards allow the analysis to run faster, as the loads are not updated with each design 

cycle update. 
174 

Case Description Mach Q NX NY NZ 

Roll 

Rate Pressure 

Inertial 

Load (G) 

1 

2.5G Pull-Up @ M 0.836 

x 1.15 0.961 4.373 0.0 0.0 2.5 0.0 5.52 N/A 

2 

1.2G Pull-Up @ M1.600 

x 1.15 1.84 6.47 0.0 0.0 1.2 0.0 9.55 N/A 

3 

-0.5G Push Over @ M 

0.836 x 1.15 0.961 4.343 0.0 0.0 

-

0.5 0.0 5.52 N/A 

4 

0.0G Push Over @ M 

1.600 x 1.15 1.84 6.47 0.0 0.0 0.0 0.0 9.55 N/A 

5 25 deg/sec @ M 0.836 0.836 3.306 0.0 0.0 0.0 0.436 5.52 N/A 

6 3-point landing 0.235 0.568 0.0 0.0 1.0 0.0 0.00 -1.6 

7 

0.5G Lateral Ground 

Maneuver 0.0 0.0 0.0 0.5 0.0 0.0 0.00 1.0 

8 

1.5G Pitch Up @ M 1.600 

x 1.15 1.84 6.47 0.0 0.0 1.5 0.0 9.55 N/A 

9 Overpressure 0.0 0.0 0.0 0.0 0.0 0.0 18.29 -1.0 

10 25 deg/sec @ M 0.836 0.836 3.306 0.0 0.0 1.6 0.436 5.52 N/A 

Sizing Results 

Parametric Weight FEM Weight Difference 

Fuselage 6,050 11,9365,886

Total Wing 22,260 35,39013,130

Tail & HPOD 3,120 5,3072,187

Total 31,430 52,63321,203

The final sized FEM yield a weight of 52,633 pounds, as shown in table above. This is 21,203 pounds 

heavier than the value predicted by the parametric weight tool. There are a number of reasons why the FEM 

weight is heavier than the parametric weight: 

1. The parametric tool is based on empirical data of previous supersonic vehicles. These vehicles did 

not have configuration similar to this 076E airplane, which has thin aft deck supporting large tail 

loads and heavy engines. 

2. The airplane has a 10% static margin, and a short moment arm for pitch maneuver. This means the 

tail will need to generate large forces in order to achieve the 2.5 G maneuvers, which means 

significant amount of structure is needed in the aft deck to support this large load. 

3. The limited volume in the wing and the aft deck does not provide efficient load path. In order to 

carry the loads, more structures are required which yield heavier weight. 

The structure team believed that the airplane structure has yet to achieve its minimal weight, and is 

optimistic that more weight can be removed. However, this will require more iteration in structural layout 

and interactions with the aerodynamic team to improve the configuration of the airplane. Here are the 

recommendations to improve the airplane’s structure and to reduce the structure weight: 

1. Increase the depth of the aft deck to allow more load bearing structures for the engine and tail 

2. Increase the depth of the inboard and outboard wing joint to allow higher bending moment of 

inertia 

3. Improve the structural layout so more spars are used to carry the load from outboard wing to the 

inboard wing box 

4. Incorporate keel beam in the fuselage to carry load from fuselage bending 

5. Conduct trade study to determine if usage of aluminum yields lighter weight 

 
175 

Conclusion from Sizing Analysis 

A comparison of the running loads analysis between this latest sized FEM and the previous iteration shows 

a significant reduction in the loads. This is the results of the updated structural layout as well as the updated 

loads criteria. Two of the critical running load cases are shown in Figures 5.1.12 and 5.1.13. 

The weight statement for the vehicle has also been updated, in Table 5.1.1, with the data from the sizing 

analysis. The MTOGW of the vehicle is still constrained at 180,000 pound, due to the requirement for sonic 

boom performance. Because of this, it was necessary to lower the fuel volume in order to counter the 

increase in structural weight, and achieve the target weight. 

Although it is possible to reduce the structural weight further to regain some of the fuel volume, using the 

recommendations in the previous section, it is possible that this airplane cannot achieve the range and 

payload requirement as an 180,000 pound airplane. It is likely that MTOGW will need to be increased to 

accommodate the structure that will be necessary to meet performance and certification requirements. 

Therefore, to be able to close the design and to have a valid finite element model for further structural 

analyses, this airplane will need to be iterated among the disciplines within the team to ensure the final 

vehicle configuration can and will address all performance requirements. Until then, this finite element 

model will serve as a good lead-in to the next design iteration, but is likely not valid for further structural 

analyses. 

Other Analyses 

The original plan called for completion of a flutter analysis, gust loads assessment and development of state 

space models. These analyses have not been completed based upon the maturity of the FEM relative to the 

overall design and performance requirements. 

5.2 Propulsion System Analysis (Option 2) 

5.2.1 WATE ++ 

5.2.1.1 - Limitations of WATE++ 

While designing the flowpath Georgia Tech ran into some difficulty using WATE++ to draw a reasonable 

flowpath because the tool is designed to provide a weight estimate, not to create a realistic flowpath. 

Inherent limitations such as restricted compressor and turbine geometries, the inability to specify individual 

stage loadings, the restriction of constant hub, tip, or meanline designs, and other factors meant that GT had 

to create several design rules which allowed GT to create a feasible flowpath from WATE++ on the first 

pass which could then be tuned with input from Rolls Royce in order to arrive at the final design. 

5.2.1.2 - WATE++ Flowpath Design Rules 

All of the design rules were implemented into NPSS/WATE++ using the built in solver functionality. The 

aeromechanical design point for the WATE++ model was set as the top of climb point; however, WATE++ 

records maximum pressures and temperatures over the entire flight envelope to calculate duct thickness and 

other parameters that depend on upper limits. 

Fan and HPC Tip Speeds 

The Fan and HPC tip speeds are set as a function of the first stage pressure ratio. Since the shaft speed 

helps to determine the turbine radial location it was important to give the designer control over this 

parameter. By relating the blade tip speed to pressure ratio it allows the designer to make intelligent choices 

that scale with engine size. The curve, shown in Figure 5.2.1, used to determine the fan tip speed was taken 
176 

without modification from “Preliminary Estimation of Engine Gas-Flow-Path Size and Weight” by Sanghi. 

The HPC tip speed was created by supplementing the fan tip speed curve with data in Hill & Peterson. 

Fan Inlet Mach Number 

The thermal cycle analysis executed by NPSS calculates a fan diameter based on the SLS mass flow, an 

assumed fan specific flow of 40 lbm/ft2

, and a hub to tip ratio listed in Table 5.2.1. The engine-vehicle 

optimization performed for the 076E made use of the fan diameter calculated in the thermal cycle analysis. 

Therefore it is important to keep the fan hub to tip ratio and specific flow in the WATE++ model equal to 

the ones specified in the NPSS thermodynamic cycle. The fan inlet Mach number is varied to maintain the 

specified fan specific flow. 

HPC / HPT / LPT Inlet Mach Numbers 

Average stage flow coefficients, listed in Table 5.2.1 for the HPC and turbines, were set by varying the 

inlet Mach number until Equation 5.2.1 is satisfied. The rotational speed falls out from the tip speed set for 

the fan or HPC. The turbine radius is calculated by WATE++ using the turbine loading parameter and the 

enthalpy rise per stage. 

( ) mid

ax

r

C

ω

φ =

Equation 5.2.1 Flow Coefficient 

Turbine Loading 

WATE++ allows the user to specify a turbine loading parameter. In order to simplify the calculations a 

constant mean line turbine with symmetrical velocity diagrams is assumed. While this would be a highly 

unlikely result for the actual design, it serves to provide a simple and quick method of relating the turbine 

loading to other input parameters. According to Analysis of Fan-Turbine Efficiency Characteristics in 

Terms of Size and Stage Number by Warner L. Stewart, the flow coefficient can be related to the turbine 

loading parameter and stator inlet angle, α1, by Equation 5.2.2. By using this relationship the designer only 

has to choose a flow coefficient and stator inlet angle. Table 5.2.1 shows the final stator inlet angles. 

( ) 1 cot

2

1

α λ

λ ⎟

⎠

⎞ ⎜

⎝

⎛ + = U

Cax

Equation 5.2.2 Turbine Loading Parameter / Flow Coefficient Relationship 

In order to keep the turbine loading within reasonable values a Smith chart is used to constrain the 

maximum possible turbine loading for the initial design pass. A Smith chart, shown in Figure 5.2.2, relates 

zero clearance turbine efficiency to the loading parameter and flow coefficient. The Smith chart is not 

corrected for tip clearance because it is assumed there will be active clearance control that greatly reduces 

tip losses. Furthermore, the actual efficiency values on the Smith chart come from test data, so they do not 

represent possible advances in technology. 

Specific Speed 

Specific speed is not used as an input into the MFTF WATE++ model; instead it serves as a sanity check 

on the combination of flow coefficients, first stage pressure ratios (i.e., rpm), and other inputs. Specific 

speed, as defined in Equation allows the designer to determine whether or not the component is designed 

in a way that will allow it to operate in an efficient manner. Curves corresponding to axial flow machines in 

Figure 5.2.3 and Figure 5.2.4 from “Introduction to Turbomachinery” by Japikse were used to determine 

whether the specific speeds that resulted from other WATE++ inputs were within reason. 
177 

( )0.75 h

m

n

shaft

s

Δ = ρ ω &

Equation 5.2.3 Definition of Specific Speed 

Stage Count Calculation 

It was found that WATE++ had very poor numerical convergence properties if the stage count was not 

specified for each component. To overcome this, WATE++ is provided with the stage count resulting from 

for each component using the simplified weight estimation routine. 

WATE++ Material Selection 

The built in WATE++ material database does not include the newest high temperature alloys or the 

advanced materials that are anticipated in 2025. Georgia Tech used appropriate materials for the various 

components, but the volume factors for each component were adjusted until the overall weight of each 

component generally agreed with the Rolls Royce simplified weight estimation method based on flow 

scaling. Rolls Royce believes the weights provided using their method are more representative of future 

material technologies. The materials used in the WATE++ model are listed in Table 5.2.2. Aspect ratio 

distributions were held constant and are listed in Table 5.2.3. 

5.2.1.3 - Georgia Tech MFTF Flowpath for 076E 

Using the design rules established for WATE++ Georgia Tech proceeded to design a flowpath for the cycle 

selected for the 076E configuration. Inputs were varied until a flowpath that did not violate any intrinsic 

mechanical constraints resulted. The final flowpath is shown in Figure 5.2.5. Pertinent input and outputs 

from the WATE++ model are summarized in Table 5.2.1. It should be noted that some of these parameters, 

such as the LPT flow coefficient, are outside of the ranges of normal design. This is due to the limitations 

of WATE++ and the inability to custom-tailor stage loading and radii distributions. The definition of the 

loading parameter used in Table 5.2.1 is shown in Equation 5.2.4. Table 5.2.4 lists the total engine 

dimensions and weight. 

2 U

Δh

ψ =

Equation 5.2.4 Definition of Turbine Loading Parameter 

5.2.2 – Flowpath 

Starting with the WATE++ flowpath defined by Georgia Tech and discussed above, Rolls-Royce used 

design architectures drawn from several Rolls-Royce engines to develop the engine cross section shown in 

Figure 5.2.6. 

The conventional combustor section and two stage turbine architecture shown in the WATE++ flowpath 

are similar in concept to the AE3007 engine. The AE3007 is a civil aircraft engine currently powering the 

Embraer 145 among other aircraft. 

The fan drive turbine of the WATE++ flowpath appeared to be most similar to fan drive turbines used to 

drive the large high bypass ratio fans in the Rolls-Royce family of engines. This design architecture was 

modeled to develop the fan drive turbine for the drawing. 

The process used in developing the general arrangement drawing was to scope out the space allocated by 

WATE++ flowpath (axial and radial) for each component. Then each component was scaled to match the 
178 

WATE++ flowpath retaining mechanical details where possible. Finally, the similarity to the “host” 

engines was modified to merge the final engine cross section into one design concept. Different challenges 

were encountered in each area of adaptation and will be dealt with in turn. 

Fan Scaling 

Even though similar Rolls-Royce engines feature a two stage fan of similar stage pressure ratios as the 

target engine, in attempting to scale the components to the WATE++ Flowpath, the Rolls-Royce design 

team found that the WATE++ airfoil aspect ratios were necessarily higher than other engines. This 

compressed the geometry somewhat in the axial direction. For the fan flowpath, this compression and the 

resulting aspect ratios were within Rolls-Royce design practice. However high aspect ratio blading at these 

stage pressure ratios is sometimes difficult to achieve meeting all aero elastic criteria normally associated 

with fan blade design. Further iterations towards lower aspect ratio blading would help the construction. 

Fan to HPC Swan Neck Scaling 

The “swan neck” space claim from the WATE++ flowpath was found to be acceptable, but the shape as 

presented was more severe than can be accomplished with current (known) aerodynamics. The issue is 

flow separation from the duct inner wall as it descends to the HPC. In order to develop the general 

arrangement drawing, a duct from a similar sized engine was scaled to the WATE ++ flowpath resulting in 

the duct shape shown. The Rolls-Royce team was able to find a swan neck duct that fit the space and 

appears to be workable. However, by 2025, flow control on the surfaces may result in shorter duct lengths 

and/or more severe duct wall angles which would favor the WATE++ shape duct. 

High Pressure Compressor Scaling 

The 11 stage compressor as defined by the WATE++ space claim (axial length) requires blade and vane 

aspect ratios higher than those associated with today’s proprietary highly loaded compressors. However, 

the blade and vane aspect ratios from the WATE++ flowpath are quite similar to previous generation 

compressor flowpaths. As the technology has improved to permit higher loading per stage, the blade and 

vane airfoils have progressively become lower in aspect ratio. The relationship is not quite as linear with 

loading as this discussion implies, but the trend is certainly true. 

Combustor and High Pressure Turbine Scaling 

Rolls-Royce aerodynamic components were found to be quite compatible with the space claim of the 

WATE++ flowpath when scaled to the engine size by flow. There was little compromise required to fit the 

geometry into the flowpath in this area. Looking forward, future turbines will almost certainly produce 

more work at reasonable efficiency. The HP spool may be driven by a single stage. It probably makes 

sense to limit the HP compressor work to that compatible with the best available single stage turbine 

technology. 

Inter-turbine Transition and Low Pressure Turbine Scaling 

The maximum flow angle found to be acceptable at the exit of the HPT (shrouded blade geometry) is 12o. 

This angle was adopted together with a shrouded second stage blade to establish the HPT to LPT turbine 

transition duct which was compatible with the space claim shown in the WATE++ flowpath. This provided 

sufficient room to include a structure through the interturbine duct so that an aft bearing could be included 

for the aft end of the HP shaft as well as a forward bearing for the LPT rotor modeling other Rolls-Royce 

engine architecture in this area. 

The Rolls-Royce team elected to “straddle” mount the LPT with both a forward and aft bearing. This 

permits a smaller fan drive shaft diameter than might otherwise be considered. An aft structure was added 

to support the LPT aft bearing, again modeling similar Rolls-Royce architectures. 
179 

Engine Flowpath 

The resulting engine meets the WATE++ overall dimensions in terms of length and diameter as shown in 

Figure 5.2.7. 
180 

5.2.3 Inlet & Nozzle Flowpath 

The nacelle outer mold lines for the 076E configuration were fixed prior to selecting the engine cycle, as 

described in Sections 3. Because of this, some inlet and nozzle design features and flexibility were 

constrained. A single pass through the design process for the inlet and nozzle flow path was performed. 

As a result, some concerns and potential short coming with the nacelle configuration selected exists and 

will be identified in this section. 

Figure 5.2.8 shows the inlet and nozzle flowpath for the selected 75 inch diameter engine for the 076E 

aircraft. The inlet is described as 3-D external compression configuration. It incorporates isentropic ramp 

to provide “nearly ideal” free-stream compression prior to throat shock. It is assumed that the baseline 

configuration will incorporates flow control technologies for optimal shock placement and stability at both 

cruise and off-design points. The take-off and low speed airflow demand will be accommodated via 

auxiliary inlet apertures located downstream of throat as shown in Figure 5.2.9. Inlet performance will be 

high at on-design flight conditions, high power, low speed with aux inlet open and cruise condition. The 

mechanical integration of aux inlet should be straightforward. The short diffuser length will require some 

form of flow control devices (e.g., micro vanes) to reduce potential for shock induced separation and 

energize diffuser flows, but may not satisfy performance and operability targets. In addition, the fixed 

ramp, bleedless inlet may not provide adequate shock control to satisfy traditional engine operability 

demands across flight envelope and power settings. 

The nozzle is incorporated into the top mounted nacelle arrangement with variable nozzle area ratio 

(a8/a9). The nozzle is an internal 2D-CD configuration with an integrated thrust reverser concept. There are 

volume reserves in the nacelle for secondary flow geometry that could be used for reducing jet noise as 

shown in Figure 5.2.10. The design assumes the baseline configuration will incorporates flow control 

technologies for matching engine exit stream with secondary airflow. At target type trust reverser concept 

is shown in Figure 5.2.11. The nacelle design and configuration constraints create some integration 

concerns. The fixed external aft nacelle compromises off design nozzle performance. The scarfed nozzle 

exit plane results in suboptimal nozzle vector, impacting both performance and aircraft control. The ultra 

short nozzle duct, particularly on top, provides little length to mix exhaust and turn flow axially; 

performance and discharge characteristics may be compromised. The upper nozzle flap is not long enough 

to completely block flow for reverse operation. In addition, secondary flow / bypass and thrust reverser 

share deployment mechanisms, complicating failure mode management. The scarf nozzle design will 

require high temperature sidewall / throat seals to minimize leakage and temperature damage to nearby 

aircraft structure. 

5.2.4 Jet Noise Certification Assessment 

Introduction 

At the outset of this contract it was realized that the scope of the noise certification assessment should be 

limited to the conceptual detail provided by the aircraft and engine performance databases. Therefore we 

selected limiting the noise prediction to screening the jet noise component at the three noise certification 

reference conditions (i.e. flyover, lateral and approach). We reached a general consensus between the 

airframe and engine team members that the most difficult noise goal condition to meet would be the lateral 

condition during the takeoff condition to meet takeoff field length and aircraft range requirements. 

The lateral condition would be dominated by the jet noise component from the engine exhaust and 

therefore we selected a jet noise exhaust velocity in the range of 1100 feet per second for the engine design 

cycle based on knowledge of past subsonic and supersonic aircraft noise certification data. Therefore the 

noise assessment would provide a comparison of the jet noise estimate to the absolute total noise goal set to 

meet the NASA N+2 contract noise certification goal range. 
181 

Supersonic N+2 Noise Goal Assumptions 

The NASA N+2 noise goal was to meet FAR Part Stage 3 minus 10 to 20 EPNdB. 

The two engine 076E configuration was sized to meet a 10,000 feet field length for an aircraft range of 

about 4000 nmi. The sizing resulted in producing a maximum takeoff weight (MTOW) = 180,000 lbs and 

maximum landing weight (MLW) = 144,000 lbs 

It was decided to initially shoot to meet Stage 3 minus 15 EPNdB (a midway position).Therefore we 

determined the absolute noise values needed to meet this requirement as follows: 

• The absolute noise levels to achieve Stage 3 minus 15 EPNdB limits would need to be 

typically: 

Flyover = 92.0 – x EPNdB 

Lateral = 97.1 – y EPNdB 

Approach = 100.7 – z EPNdB 

• The proposed initial delta goal increments relative to Stage 3 were selected as follows: 

x = 7, y = 3 and z = 5 for a total of – 15 EPNdB 

Hence the absolute goals were set at: 

Flyover = 85.0 EPNdB 

Lateral = 94.1 EPNdB 

Approach = 95.7 EPNdB 

Another benchmark in setting these absolute goals was the fact that from an operational point of view it is 

well documented that the noisiest stage 3 aircraft existing in today’s fleet (e.g. retrofitted older Stage 2 

aircraft and MD-80 types) during departure have been perceived to be as objectionable neighbors in the 

communities surrounding the airports. For example, the MD80 noise certification Lateral level for an 

engine sea level static takeoff thrust of 20,850lbs = 96.0 EPNdB. Therefore the initial goal of 94.1 EPNdB 

for the 96E configuration is in the favorable lower noise direction. 

Noise Certification Measurement Reference Locations 

The FAR Part 36 reference noise certification measurement locations are illustrated in Figure 5.2.12 for the 

Flyover, Lateral and Approach conditions. 

Noise Screening Assumptions 

Unsuppressed jet noise was estimated at FAR Part 36 measuring locations for the following reference day 

conditions: 

ISA+10o C day 

70% Relative Humidity 

Zero Wind 

Sea Level 

The plan was to assess the need for jet noise suppression devices to meet total noise goals including 

margins available for the contribution of fan (forward & aft) , turbine, core and airframe noise sources form 

the unsuppressed jet noise screening results 

Jet Noise Prediction Assumptions 

We used the ANNOP jet noise prediction code for a mixed flow turbofan cycle with internal mixer that 

uses the mass weighted jet velocity of the fan and core flows at the nozzle exhaust exit station. Additionally 

we assumed the following: 

Assumed fully mixed flow at nozzle exit (i.e. No hot spots) 

Assumed axi-symmetric nozzle exit 
182 

No account taken for potential self noise contribution of mixer 

No account taken for any possible shielding of distributed jet noise sources by twin verticals 

(lateral) and aft deck platform (flyover) 

Jet Noise Prediction Results 

The following jet noise prediction results were obtained for the 076E configuration 

FAR Part 36 Reference Conditions - Ground Rules 

The ground rules for the three reference conditions were as follows: 

a) Flyover Conditions 

- Monitor point at 21325 ft (6500m)from brakes release 

- MTOW 

- Maximum takeoff power (de-rated) from brakes release to meet TOFL requirements 

- V2+10 TAS climb-out speed 

- First screening without cutback procedure 

- If needed thrust cutback at approx 4000 feet before monitor to engine power to maintain 

level flight with one engine inoperative or 4% gradient power whichever is higher 

- Assume an instantaneous cutback flight profile 

b) Lateral Conditions 

- MTOW- 

- Maximum Takeoff power (de-rated) to meet TOFL requirements 

- V2+10 TAS climb-out speed 

- Noise monitor at 1476 feet (450m) to the side of the runway extended centerline 

- Assume peak lateral noise occurs at 1000 feet AGL. This represents an angle of elevation 

from monitor to aircraft of 34.1o (Slant range = 1782 feet) 

- Assume lateral attenuation from ANOPP 

- No impact of cutback procedure on peak Lateral noise 

-

c) Approach Conditions 

- MLW 

- Approach monitor at 6560 ft (2000m) from runway threshold (394 feet) 

- Aircraft to maintain -3 degree glide-slope 

- Define noisiest configuration - dirtiest aerodynamic - highest drag condition (30 deg flap) 

- Front and main gear down 

- Low speed high lift devices out 

- Vref+ 10 TAS approach speed 

And the results are: 

Flyover (No cutback) 

Altitude = ~ 2800 feet 

V2+10 True Airspeed = 201 knots 

Total Net Thrust = 39,588lbs 

Vj = 1235 fps 

Predicted Jet Noise Level = 87.3 EPNdB 

Total Noise Level Goal = 85.0 EPNdB 

Difference relative to goal = +2.3 EPNdB 

Lateral (1476 ft from runway centerline) 

Altitude = 1000 feet 

V2+10 True Airspeed = 201 knots 

Total Net Thrust = 42,522 lbs 

Vj = 1238 fps 
183 

Predicted Jet Noise Level = 91.7 EPNdB 

Total Noise Level Goal = 94.1 EPNdB 

Difference relative to goal = -2.4 EPNdB 

Approach 

Altitude = 394 feet 

Vref+10 True Airspeed = 165 knots 

Total Net Thrust = 26,538 lbs 

Vj = 979 fps 

Predicted Jet Noise Level = 93.6 EPNdB 

Total Noise Level Goal = 95.7 EPNdB 

Difference relative to goal = -2.1 EPNdB 

Jet Noise Screening Conclusions 

Based on the jet noise screening results we can draw the following conclusions: 

1. Engine cycle (Vj = 1238 fps) at de-rated power adequate to achieve lateral total noise goal ( 

jet noise -2.4 EPNdB below goal). Below benchmark MD-80 noise certification level 

2. Flyover Jet noise (without cutback) needs to be further reduced to achieve a 

reasonable margin relative to total noise goal (current jet noise at 2.3 EPNdB above goal) 

3. A larger approach jet noise margin relative to total noise goal needed to account for other 

noise sources (fan, turbine, core & airframe) contributing to total noise (current jet noise – 1.9 

EPNdB below goal) 

4. Stage 3 – 15 EPNdB should be doable 

5. Stage 3 – 20 EPNdB (i.e. to meet probable Stage 5 limits in 2016 will be a stretch) 

5.2.5 Derated Takeoff and Jet Velocity Relationship 

The NPSS parametric engine model and Model Center optimization focused on maximizing fuel efficiency 

while meeting the community noise goal. To do this, mixed jet velocity is used as surrogate for community 

noise. A target jet velocity of 1100 ft/sec was set. Derated takeoff thrust capability and Automatic Takeoff 

Thrust Control System (ATTCS) FAR Part 25 regulations were used to minimize takeoff jet velocity. The 

transonic pinch point and the top of climb flight conditions are engine sizing conditions. To take advantage 

of this relationship, Logic was added to the NPSS model to allow input of jet velocity which set the derated 

takeoff thrust rating. Figure 5.2.13 shows that a 21% lower jet velocity, at lateral and flyover conditions, 

are obtainable relative to the full thrust capability of the engine. Also, shown on the figure is the 39% 

derate available while still meeting the 10,000 ft balanced field distance. 
184 

Tables and Figures for Section 5.0 

Figure 5.1.1 Flight Envelope for Design Loads Development. 

Figure 5.1.2 Design Load Conditions 
185 

Figure 5.1.3. Aft Deck and Tail Joint\ 

Figure 5.1.4. Inboard and Outboard Wing Joint. 
186 

Figure 5.1.5. Wing & tail skin materials. 

Figure 5.1.6. Internal structure materials.
187 

Figure 5.1.7 076E Structural Layout. 

 
188 

Figure 5.1.8. Streamwise Running Load Plot, 1.5G Pull-Up at Supersonic Dive Speed. 

Figure 5.1.9. Spanwise Running Load Plot, 40 degree/sec Roll at Subsonic Speed.
189 

Figure 5.1.10 076E Keel Structure. 

Figure 5.1.11. 076E Structural Analysis Process. 
190 

Figure 5.1.12 Comparison of Streamwise Running Loads during supersonic pull-up maneuver. 

Figure 5.1.13 Comparison of Spanwise Running Loads during subsonic roll maneuver. 
191 

Table 5.1.1 Weight Statement Update using Sized Structural Weight. 

Figure 5.2.1 Fan and HPC Tip Speeds vs. First Stage PR 

Blade Tip Speed Estimation

y = -329.24x2 + 2239x - 1307

y = -329.24x2

 + 2239x - 1157

0

500

1000

1500

2000

2500

1 1.2 1.4 1.6 1.8 2

Fan

HPC

Poly. (Fan)

Poly. (HPC)
192 

Table 5.2.1 WATE++ Component Information 

 Fan HPC HPT LPT 

Pressure/Expansion Ratio 2.146 16.821 4.24 3.56 

Hub to Tip Ratio 0.35 0.68 0.905 0.87 

Stage Count 2 11 2 4 

Rotor Speed (RPM) 4476 9730 9730 4476 

Design Flow Coefficient 0.638 0.81 0.35 1.2 

Stage Loading Parameter (per stage) - - 2.25 2.93 

First Stage Stator Inlet Angle 67o 74o 

Corrected Tip Speed (ft/s) 1394 1168 - - 

Component Weight (lbm) 3154 1040 684 1211 

Hub to Tip Ratio 0.35 0.681 0.91 0.87 

Length (in) 44.66 25.32 6.76 15.73 

 

Figure 5.2.2 Smith Chart 
193 

Figure 5.2.3 Compressor Efficiency vs. Specific Speed 

Figure 5.2.4 Turbine Efficiency vs. Specific Speed 
194 

Table 5.2.2 WATE++ Material Selection 

Component Blade Material Disk Material

Fan Ti-6Al-2Sn-4Zr-2Mo Inconel 718

HPC

If(max Temp < 1250oR)

Ti-6Al-2Sn-4Zr-2Mo

If(max Temp > 1250oR)

Inconel 718

Inconel 718

HPT MAR-M247 Udimet 700

LPT Udimet 700 Rene 95

Table 5.2.3 WATE++ Aspect Ratio Distributions 

Component First stage Aspect Ratio Last Stage Aspect ratio

Fan 3.2 2.7

HPC 2.5 1.5

HPT 1.25 2.0

LPT 3.3 5.3

Figure 5.2.5 GT WATE++ Flowpath for MFTF 
195 

Table 5.2.4 External Engine Dimensions 

Total Weight (lbm) 7523 

Total Length (in) 225 

Total Diameter (in) 83 

Figure 5.2.6 Engine General Arrangement developed by Rolls-Royce from WATE++ flowpath and cycle 

information provided 
196 

Figure 5.2.7 The R-R General Arrangement Matches the WATE++ Flowpath 

Figure 5.2.8 Inlet and Nozzle Flowpath with Engine Fan Diameter of 75 inches 
197 

Figure 5.2.9 076E Inlet with Auxiliary Inlet 

Figure 5.2.10 076E Nozzle w/ Secondary Flow Geometry 
198 

Figure 5.2.11 076E Nozzle w/Thrust Reverser Deployed 

Figure 5.2.12 - Noise Measurement Reference Locations 

 
199 

Figure 5.2.13 Benefit of Derated Takeoff On Jet Velocity 

NPSS BN2-GR-003.B / -076E / Max Range Opt Cycle

Lateral & Flyover Conditions, 201 KTAS (0.31 Mn), ISA+10C

0

5000

10000

15000

20000

25000

30000

35000

40000

0 1000 2000 3000 4000 5000 6000

Altitude, feet

Net Thrust, Lbs

1000

1200

1400

1600

1800

2000

2200

2400

2600

Mixed Jet Velocity, ft/sec

All-Engine T/O Rating (Fn)

Engine-Out T/O Rating (Fn)

100% Fn

Lateral Noise Pt Fn

Acoustice Ref Pt Fn

All-EngineT/O Raitng (Vj)

Engine-Out T/O Rating (Vj)

100% Vj

Lateral Noise Pt Vj

Acoustic Ref Pt Vj

39% 45%

21%
200 

6.0 Technical Shortcomings and Portfolio Gap Engineering Plan for 

Continued Configuration Development 

The N+2 time-frame allows approximately 8 years from today for technology, configuration, and design 

process development before committing to the available and near-term forecast levels. While any 

technology or configuration feature with a credible promise of reducing noise, reducing boom, raising fuel 

efficiency or lowering weight could help achieve the fuel efficiency, emissions and boom goals, a few 

stand out as critical. 

Summary of key enablers: 

- Aero- Propulso- Servo- Elastic (APSE) design and evolved fly-by-wire, multi-function surfaces, 

and multi-input/multi-output (MIMO) design. 

 - Configuration features that enable both low boom and low drag. 

 - Configuration optimization via CFD & FEM, MDA/O tools. 

 - Inlets & nozzles that are simple, efficient, and low noise. 

 - Durable engine materials, reduced SFC optimized engine cycles. 

 - Design for low speed performance, handling, field performance, noise. 

 - Credible & achievable certification requirements (boom, loads, emissions, etc.) 

A more detailed compilation identifying specific technical shortcomings and portfolio gaps presently facing 

the N+2 configurations was prepared and categorized. They are in priority order under each Category. 

Multi-Disciplinary Optimization and Trades: (In priority order)

– Outer-loop MDAO, especially between aero & structures, needs multiple cycles. N+2 

aero-structural iteration not converged. 

– Propulsion system sizing and technology selection. 

– Overall thrust & wing area sizing along with technology selection. 

Airframe/Structures (In priority order)

– Light weight detail design concepts to meet weight goals. 

– Structure layout & materials systems needs more development. 

– Further optimization work on the 076E structural configuration exploring whether global 

optimum has been reached. 

– Certification i.e. loads criteria, rotor burst. 

– Systems packaging and accessibility. 

– Flight deck visibility and electronic supplemental vision. 

– Certifiable fuel/CG management system. 

Aero (In priority order)

– NASA 65 to 70 PLdB goal not achievable with current configuration & length. 

– Non-linear boom optimization to achieve goal PLdB, i.e 100 vs 90 while maintaining 

L/D. 

– Improved target wave forms compatible with configuration integration and certification. 

– Low speed performance and handling qualities risks. 

– Takeoff field performance margins with derated thrust. 

– Propulsion system external flow aero interference & propulsion airframe integration. 

Propulsion System (In priority order)

– Inlet/Diffuser technology opportunities 

• Fixed Ramp, Bleedless Inlet: Simplifies inlet by eliminating porous surfaces, 

bypassing & momentum loss. But may not provide adequate shock control to 
201 

satisfy traditional engine operability demands across flight envelope and power 

settings. 

• Flow Control Devices: (e.g., Micro Vanes) Will Reduce Potential for Shock 

Induced Separation and Energize Diffuser Flows. But may not satisfy 

performance and operability targets. 

– Nozzle technology opportunities (in priority order): 

• Scarfed Nozzle Exit Plane: Facilitates engine integration, but results in 

suboptimal nozzle vector, impacting both performance and aircraft control. 

• Fixed External Nacelle Exit: Compromises off design nozzle performance. 

• Ultra Short Nozzle Duct: (Particularly on top) Provides little length to mix 

exhaust and turn flow axially, so performance and discharge characteristics may 

be compromised. 

– N+2 Engine size optimization opportunity: 

• Engine / Nacelle Diameter Target vs Status, (66 vs 75 in.), High jet velocity at 

target diameter, aggravates inlet & nozzle length shortcomings. 

The investigation of the propulsion system revealed several other specific technologies and opportunities 

for development that are relevant to any continued work on N+2-type airplanes. Areas recommended for 

continued development work include: 

Propulsion System Roadmaps - Outline 

– Upper Inlet - Compact Inlet designs with high pressure recovery and distortion tolerance 

• Establish key design requirements and objectives 

• Active and passive flow control techniques 

• Min L/D diffuser design studies 

– High performance lightweight 2-D scarf nozzle designs 

• Establish key design requirements and objectives 

• Aero/mechanical design and analysis trades 

– Reduced weight, high performance mixer / ejector nozzle 

• Establish target and status technology trends 

• Physical design constraints and features 

– Higher fidelity engine sizing and cycle optimization 

• Improved component and subsystem performance models 

• Mixer/ejector trades and MDAO at the vehicle level 

Engine Sizing and Optimization (Propulsion priority 1) 

– Resize engine and optimize cycle with updated vehicle configuration 

• Use engine company proprietary parametric models with higher performance 

and additional N+2 engine technologies 

• Include additional cycle / architecture design variables 

Mixer / Ejector Trade Studies / MDAO (Propulsion priority 2) 

– Use parametric model with M/E performance logic 

• Establish M/E vs engine size and cycle design space 

• Development technology goals / targets required for M/E to buy its way onto the 

vehicle (weight, thrust loss, Vj reduction/ effectiveness) 

Exetrnal Compressible Inlet (Propulsion priority 3) 

– Compact inlet design with high pressure recovery and distortion tolerance 

• Passive and active control. 

• Minimum length diffuser design. 

(Propulsion technology shortcomings & gaps, continued) 

 Upper Inlet Design Refinement (Propulsion priority 4) 

– 3-D RANS Assessment of Main and Aux Inlet Performance at Critical Flight Points; 

Refine Lines to Achieve Target Performance 
202 

 2-D Scarf Nozzle Design Refinement (Propulsion priority 5) 

– Define Changes to Nozzle Envelope In Order to Close on Concept 

– 3-D RANS Assessment of Low Speed Operation with Base Relief and Main Nozzle 

Performance at Critical Points in Flight Envelope 

– Refine Nozzle Lines / Mechanical Concept of Operation and Verify Performance Targets 

Are Achievable 

 Acoustics—Jet Noise (Propulsion priority 6) 

– Flyover – Develop and optimize a cutback procedure 

– Flyover – Assess Program Lapse Rate (PLR) procedures to improve flyover altitude 

– Flyover & Lateral – Consider nozzle noise suppression devices 

– Approach - Improve low speed aerodynamic devices (increase L/D i.e. less thrust/rpm) 

 Georgia Insitute of Technology and Rolls Royce technology shortcomings and portfolio gaps 

are shown in Appendix 3.0. 
203 

7.0 Summary and Conclusions 

7.1 Configuration Deliverables 

Non-proprietary conceptual designs for supersonic airliners two technology generations into the future 

(N+2, 2020-2025 EIS) have been developed: the 100 passenger model 765-072B, and the 30 passenger 

model 765-076E. Both have the potential to achieve a trans-Atlantic range of 4000nm. Both of these 

“concept planes” are described in sufficient detail for them to serve as platforms for NASA technology 

assessments and as points of departure for other future work. 

3D CAD lofts of the outer mold lines (OML) of these two configurations, along with selected sample 

geometries of other configurations considered, are available for use in future Multi-Disciplinary Analysis 

and Optimization (MDAO) studies. 

7.2 Reference Vehicle Requirements 

Aircraft concept development for the N+2 study was guided by first establishing a reference set of vehicle 

requirements appropriate for a commercial supersonic airliner hypothetically entering passenger service as 

early as the years 2020-2025. Important factors that could be expected to drive the design, certification, and 

economic viability of such an aircraft were identified. These factors included the long-standing supersonic 

airliner challenges of flight safety, low fuel-burn (for both operating economics and environmental 

acceptability), meeting ever more stringent takeoff and landing noise restrictions, minimizing sonic boom, 

low emissions engine combustion, engine and airframe durability, and fitting into any of the world’s air 

traffic control regimes with minimal disruption. It was concluded that environmental concerns would be 

most critical in that regardless of vehicle payload, price, or performance, no future supersonic aircraft can 

be considered “viable” if it is perceived as an outlier raising environmental concerns disproportionate to 

any potential economic or social benefits. The combination of requirements was determined to be best 

satisfied by configurations flying slower than Concorde (Mach 1.6-1.8) at a maximum altitude of 55,000 

feet, employing moderate bypass ratio mixed flow turbofan engines potentially having some variable cycle 

features, with variable area nozzles, and mixed composite-metallic airframe construction. 

7.3 Design Space Exploration And Trades 

Multiple aircraft general arrangements were investigated to meet the identified requirements at an initial 

conceptual design level. Initial TOGW and engine sizing were determined using proprietary engine 

definitions from previous Boeing internal studies, with appropriate scaling and technology assumption 

adjustments. A large design space was investigated around selected configuration “species” using several 

MDA/MDO approaches. Trade sensitivities for passenger payload, design range, cruise Mach, cruise 

altitude, and cabin shape constraints were conducted with low fuel burn or low sonic boom as alternate 

figures of merit. 

A sophisticated non-proprietary parametric turbofan engine model was created for MDA/MDO studies, and 

was exercised extensively in a set of airframe-to-engine cycle matching trades and design space 

sensitivities. This engine model, based on the NPSS tool, was created using engine company inputs for 

engine architecture and stage efficiency assumptions and technology level setting for the N+2 time frame. 

This tool proved invaluable in optimizing the final configuration choices and provides propulsion modeling 

capabilities considerably more advanced than those available in previous supersonic studies (e.g. HSR’s 

Design Optimization Synthesis System –DOSS). 

7.4 Down-selected Configurations and Sonic Boom Considerations 

Of various configurations studied, the first down-selected configuration, the -072B, was recommended by 

Boeing as the best representation of a low fuel burn, low emissions, higher passenger capacity 

configuration for the N+2 time frame. The -072B is intended to represent a design of challenging but 
204 

reasonable technical risk---a relatively straight-forward extension of experience obtained on the NASA 

HSR effort and the subsequent industry studies. The aircraft was configured to carry a number of 

passengers similar to Concorde at trans-Atlantic-plus ranges but at higher cabin comfort levels. The Mach 

1.6-1.8 capable airplane would also have much more efficient subsonic cruise performance, would meet the 

same community noise standards as then-year subsonic aircraft, and could exceed the N+2 fuel mileage 

goal of 3 seat-nautical miles per pound of fuel burned. The airplane size and fuel use could potentially 

allow reasonable ticket surcharges when compared to premium subsonic fares or bizjet charter rates. The 

general arrangement, wing design, etc. for this configuration, have been biased toward meeting certification 

goals and fuel mileage performance with reasonable risk---it does not incorporate low-boom shaping. 

While still having an N-wave sonic boom and flying at altitudes lower than 55,000ft, the -072B would still 

keep maximum sonic boom levels below Concorde, meeting the minimum reduction proposed for future 

over-water supersonic flights. As with other “un-shaped” boom configurations, if future flight experiments 

and advances in real-time weather monitoring, flight management systems, and “free flight” routing enable 

reliable flight at “threshold Mach”, then operation over land or in coastal regions might be allowable at up 

to Mach 1.15+/- without producing a boom at ground level. 

The -076E, initially developed as an alternate concept to illustrate the impact of low-boom specifications 

on configurations, was down-selected by NASA for further development, additional definition and 

assessment due to its focus on low sonic boom shaping as a key long-lead enabling technology. It 

represents a higher risk configuration and presents larger “technology stretch” goals for performance, 

structures, and other elements. Sonic boom mitigation was highlighted in the N+2 vehicle requirements as a 

significant hurdle to environmental acceptability of any future supersonic civil aircraft. The economic 

productivity and practicality of supersonic business jets, a likely precursor to any future generation of 

airliners, may depend on the ability of operators to fly supersonically over land on a regular basis, at least 

in designated “boom corridors”. Increased worldwide environmental awareness and concern over the 

impact of strong sonic booms will likely dictate that even an aircraft designed for supersonic operation only 

over water will need to incorporate some degree of boom “softening” to stay below the 2.2 psf 

overpressures generated by Concorde (substantially below the 3-3.5 psf forecast for the 300+ passenger 

HSCT’s studied in the 1990’s). Operation at maximum supersonic speed in remote over-land corridors and 

coastal areas maybe permitted at overpressures of 0.5-0.7 psf (PLdB’s in the high 80’s to low 90’s), while 

operations over the continental U.S., even in corridors, will likely require PLdB levels in the low 80’s. 

In recent years, significant advances have been made in the ability to define specific target disturbance 

functions for “shaped” low-boom wave forms and to select aircraft general arrangements and features that 

greatly improve the likelihood of converging on the corresponding target using CFD-based optimization 

techniques. The use of shaped wave forms provides the minimization of the largest individual shocks in the 

signature and increases shock rise times---factors which dominate the perceived noise level decibels 

(PLdB). In spite of advances in low-boom design, the physics of sonic booms dictates that the boom noise 

level produced by a given signature shape is fundamentally related to vehicle length and gross weight. The 

theoretical minimum vehicle length required to produce a shaped signature is still given by the Seebass 

(“SEEB”) formulation which is now several decades old. More sophisticated alternate disturbance 

functions or more extreme design features can produce lower PLdB levels for a given weight or provide 

added configuration integration freedom, but generally at a penalty in overall vehicle length, practicality, or 

loss of boom “robustness”. These fundamental physical limitations make it increasingly difficult to achieve 

a given goal PLdB level as the size and payload of the configuration is increased or more constraints are 

placed on the design for integration, stability and control, or structural/APSE risk management. This 

means that a goal PLdB loudness level that is fairly easily achieved with a single seat “clean sheet” 

research aircraft becomes much more difficult to achieve with a modified existing aircraft, and is much 

more difficult to achieve with an airliner sized aircraft than with a small SSBJ type. High performing, 

multi-Mach optimized Boeing SSBJ concept planes that “close” from the integration and performance 

standpoint with reasonable technical and certification risk have been developed in recent years using 

Boeing proprietary design features and processes. Such designs could potentially meet all environmental 

and certification requirements in the N+1 time-frame (year 2020), while providing “shaped” sonic boom 

PLdB levels in the low 80’s (at takeoff gross weights on the order of 100,000 lbs). Achieving a similar 

level of boom on an airliner of two or more times the weight while maintaining high performance levels is 

a formidable task, requiring the consideration of more unique general arrangements and the projection of 
205 

more advanced technologies which could be available in the years 2020-2025 and beyond. The same tools 

used to develop recent proprietary IR&D concept planes were used under the present NASA study to 

configure the -076E which is representative of such a future small low-boom airliner concept. 

7.5 Assessment of the -076E and Development/ Technology Portfolio Gaps 

Linear-based aerodynamics tools and “Quick-O” conceptual design weights were used to establish year 

2025 “goal” performance levels for the -072B and -076E. Driven by the sonic boom requirement, the -076E 

used more aggressive initial integration assumptions, and higher effective percentage of supersonic L/D 

technology assumptions than would have been used in the HSR program or more recent industry SSBJ 

studies. Linear-based aerodynamics tools were also used to establish the “theoretically achievable” sonic 

boom potential of the -076E concept. An alternate “stretched” version (-076F) was investigated as part of a 

boom sensitivity to wing aft-loading and body length, but it was dropped after consultation with NASA due 

to concerns over very high structural /APSE risk levels this might introduced into the pending FEM 

analysis. 

Because the -076E concept represents a much more aggressive “clean sheet” approach that does not relate 

well to configurations studied under HSR, nor to more recent industry proprietary studies, the status levels 

of performance, weights, internal loads, flutter risks, loadability, would be expected to be farther away 

from the configuration’s eventual potential than with a more conventional concept. A single cycle CFD 

optimization for drag level, and “one and a half” cycles of Finite Element Method analysis were conducted 

to determine the existing shortfall between the “as-drawn” status and the goal levels projected for the 

concept assuming year 2025 technology. Results from the initial CFD optimization showed that the 

projected L/D levels should be quite achievable with today’s industry tools. However it was concluded that 

retaining that performance level while simultaneously reshaping the vehicle’s contours to capture the 

identified low-boom potential represents a significant technology challenge. Improvements will be needed 

in computational efficiency, optimization algorithms, CFD accuracy, and designer insights by 2025 if such 

goals are to be achieved with any certainty. 

A partial FEM cycle was initially conducted based on a notional internal structural arrangement. This 

model was updated for the full FEM cycle with “first cut” design improvements and loads assumptions 

based on the initial results and observed behaviors that were similar to previous Boeing proprietary 

configurations. The FEM results gave important data on the running loads, load-paths, structural sizing 

challenges and identified key weight and loads-sensitive zones for this class of low-boom aircraft. These 

FEM results can be used to identify configuration concept improvements, and have highlighted the types of 

design development and technology gaps which would need to be filled by 2025 for such a concept to be 

viable. Such gaps would include improved structural arrangements, alternate materials mix, high 

strength/high modulus materials, and the need for rapid multi-cycle MDA/MDO iterations between aero 

and loads that include coupled internal and external topology changes. A number of specific candidate 

structural improvements and areas for trade-offs with aerodynamics were identified. 

The reference vehicle requirements established for the N+2 study specified that any future supersonic 

airliner would need to meet then-year community noise levels. A projection was made that the likely noise 

restrictions in the 2020+ time frame will likely require a cumulative reduction of 15+ dB relative to “Stage 

3” FAA/ICAO reference levels. The report contains a detailed discussion of the associated logic and 

requirements specific to each noise measuring point. The scope of the N+2 study and available propulsion 

system definitions precluded a detailed assessment of all of the source components of the -076E concept. 

However, as was done for the initial configuration selection studies and MDA/MDO trades and 

sensitivities, the assumption was made that jet noise would be the dominant source and engine mixed jet 

velocity (Vj) was evaluated as a surrogate for overall noise levels. A detailed analysis of the jet velocity 

noise of the -076E showed it would likely be capable of meeting the Stage 3 minus 15dB cumulative level 

with the NPSS-optimized engine cycle and a simple light weight nozzle. Turbo-machinery noise, inlet 

noise, and airframe noise components were not included, but neither was credit for optimized takeoff thrust 

lapse rate management (an advanced version of “PLR”) or any potential noise shielding effects of the V-tail 

arrangement and over-wing engines. An example hypothetical low-noise inlet and variable area thrust 

reversing nozzle concept are shown as part of the configuration definition as “technology placeholders”. 
206 

The development of high efficiency, distortion tolerant, compact inlet technology and integrated nozzle 

concepts with durable high-temperature seals were cited as additional “portfolio gaps” that would need to 

be filled to make such an airliner viable. It was further noted that if the actual then-year noise requirement 

calls for larger margins at sideline, or the net cumulative delta from Stage 3 needs to be 18-20dB ---a 

distinct possibility, then; a) other noise sources from the airframe, inlet and turbo-machinery will begin to 

intrude on the jet noise, and b) de-rate and some form of PLR may not be sufficient without some form or 

“low impact” suppressor nozzle technology. 

 
207 

8.0 References 

1. S. L. Baughcum, I. C. Plumb, I. C., and P. F. Vohralik, "Stratospheric Ozone Sensitivity to Aircraft 

Cruise Altitudes and NOx Emissions", in Proceedings of the European Conference on Aviation, 

Atmosphere and Climate, R. Sausen and G.T. Amanatidis (Eds.), Air Pollution Research Report No. 83, 

Commission of the European Communities, 197-206, 2004. 

2. D. J. Wuebbles, M. Dutta, K.O. Patten, and S. L. Baughcum, " Parametric Study of Potential Effects of 

Aircraft Emissions on Stratospheric Ozone", in Proceedings of the European Conference on Aviation, 

Atmosphere and Climate, R. Sausen and G.T. Amanatidis (Eds.), Air Pollution Research Report No. 83, 

Commission of the European Communities, 197-206, 2004. 

3. S. R. Kawa, J. G. Anderson, S. L. Baughcum, C. A. Brock, W. H. Brune, R. C. Cohen, D. E. Kinnison, 

P. A. Newman, J. M. Rodriquez, R. S. Stolarski, D. Waugh, and S. C. Wofsy, Assessment of the Effects of 

High-Speed Aircraft in the Stratosphere: 1998, NASA Technical Report TP-1999-209237, 1999. 

4. I Isaksen, C. Jackman, S. L. Baughcum, F. Dentener, W. Grose, P. Kasibhatla, D. Kinnison, M. Ko, J. 

McConnell, G. Pitari, and D. Wuebbles, "Modeling the Chemical Composition of the Future Atmosphere,", 

Chapter 4 in the Intergovernmental Panel on Climate Change (IPCC) Special Report on Aviation and the 

Global Atmosphere, [J. E. Penner, D. H. Lister, D. J. Griggs, D. J. Dokken, and M. McFarland (eds.)] 

Cambridge University Press, 1999. 
208 

Appendix 1 - Description of the 072B 

Figure A.1.1 072B General Arrangement 

 WI NG CANARD VERTI CALAREA ( sf ) 4600 210 335VOLUME - 0. 0327 0. 0389

ASPECT RATI O 2. 6 2. 71 1. 37TAPER RATI O 0. 045 0. 275 0. 281LE SWEEP ( deg) 68/ 40 40 39. 4TE SWEEP ( deg) - 15/ 6 0 0MAC ( i n) 1082 117 207XREF ( i n) 1692 592. 2 2509. 8SPAN ( f t ) 115 23 21

PI LOT EYEX = 450Y = -24Z = 29. 8

285. 95

2 3 1 0

116. 85

MAC

1099 . 71

LC

FS 1600

PI LOT EYEX = 450Y = - 24Z = 29. 8

NOTI ONAL CGX = 1692

2820

2 35

207. 45

MAC

4 x H40x 14. 5- 19 MLG TI RES 2 x H22x 8. 25- 10 NLG TI RES 352 . 20 29 4

138 1. 82

115 2

418 . 55

34 11

9 Fi r st

36-i n pi t ch

90 Economy

32-i n pi t ch

 Cl ass ( %) Cart s Cart Rat i o Lav at ory

Rat i o

At t endentRat i o

Fi r st 9. 1 3 0. 333 9 -Economy 90. 9 5 0. 055 45 -Tot al - 8 0. 081 33 25

MI NI MUM CROSS SECTI ONS

1/ 50

FS 717TY PE I

42 X 66

FS 1306

T YPE I I I

20 X 36

FS 191 7. 5

TYPE I

42 X 6 6

FS 23 

APB

FRONT PRESSURE BULKHEAD EE BAY 

FS 1600
Figure A.1.2 072B Features 

Figure A.1.3 Modifications Made Post Analysis 

Synthetic 

Vision Flight 

Deck

Full-flying 

Vertical Tail for 

SS Engine Out

Aft Stowed 

Bulk Cargo

Compromise 

Engine Meets 

1100 fps T-O

Canard 

Integrated 

Below Floor

Passenger 

Accommodations 

per 1992 IAC Rules

Area Ruled Nonconstant Crosssection Enabled by 

Composites

Synthetic 

Vision Flight 

Deck

Full-flying 

Vertical Tail for 

SS Engine Out

Aft Stowed 

Bulk Cargo

Compromise 

Engine Meets 

1100 fps T-O

Canard 

Integrated 

Below Floor

Passenger 

Accommodations 

per 1992 IAC Rules

Area Ruled Nonconstant Crosssection Enabled by 

Composites

• Resheared inboard wing (to avoid passenger compartment)

– Root airfoil likely too thick

• Reshaped fuselage for...

– ...non-circular section with flat floor

– ...integrated flight deck

– ...landing gear integration

• Lowered canard for integration with floor

• Shifted nacelle outboard for improved gear integration
210 

Figure A.1.4 072B Interior Arrangement 

Figure A.1.5 072B Landing Gear Integration 

Figure A.1.6 072B Cockpit Vision 

9 Fi r st

36- i n pi t ch

90 Economy

32- i n pi t ch

 Cl ass ( %) Car t s Car t Rat i o Lavat or y

Rat i o

At t endent

Rat i o

Fi r st 9. 1 3 0. 333 9 -

Economy 90. 9 5 0. 055 45 -

Tot al - 8 0. 081 33 25

Passengers accomodated per 1992 IAC rules, 

dual class short/medium range

9 Fi r st

36- i n pi t ch

90 Economy

32- i n pi t ch

 Cl ass ( %) Car t s Car t Rat i o Lavat or y

Rat i o

At t endent

Rat i o

Fi r st 9. 1 3 0. 333 9 -

Economy 90. 9 5 0. 055 45 -

Tot al - 8 0. 081 33 25

Passengers accomodated per 1992 IAC rules, 

dual class short/medium range

• Main Landing Gear

– Dual-tandem gear with 4 x H40x14.5-19

– Fold forward and inboard

– No unusual kinematics required

• Nose Landing Gear

– Dual with 2 x H22x8.25-10

– Fold forward

– Cantilevered off front of pressure 

bulkhead (no doghouse)

-50

-25

0

25

50

-150 -125 -100 -75 -50 -25 0 25 50 75 100 125 150

Azimuth (deg)

Elevation (deg)

Synthetic Vision will be required forward

Required Vision

MIL-STD-850

Provided Vision
211 

Figure A.1.7 072B Conceptual Structural Layout 

• Developed initial structural layout as starting 

point for FEM task

• Structure will evolve as concept analysis 

proceeds
212 

Figure A.1.8 072B Fuel Tank Layout 

 

8 659 1991

7 659 1991

6 477 1816

5 477 1816

4 225 1651

3 225 1651

2 526 1438

1 526 1438

Tank Volume (cf) X centroid

8 659 1991

7 659 1991

6 477 1816

5 477 1816

4 225 1651

3 225 1651

2 526 1438

1 526 1438

Tank Volume (cf) X centroid

Total Volume 3,400 cf

Total Fuel Weight 160,000 lb

3

1 2

4

5 6

7 8

Fuel required under 

floor in rear cabin
213 

Appendix 2 - The Boeing In-House Multidisciplinary Design and Analysis 

Tool 

Multi-Disciplinary Analysis (MDA) capability has been around since the dawn of airplane design embodied in the 

designer. Introduction of computers has enabled rapid analysis cycles in various engineering disciplines. Early 

implementation of a Multi-Disciplinary computer environment was introduced with CASES (Computer Analysis 

and Sizing Evaluation System) in 1985? Currently, Multi-Disciplinary Analysis has become highly integrated and 

automated due to faster processors, platform independent programming languages, and the speed of the internet. 

Figure A.2.1 outlines the current Multi-Disciplinary Analysis used on the Supersonic NRA N+2 Contract. This 

BR&T developed capability addresses Advanced Design engineering disciplines associated in design and analysis of 

a supersonic air vehicle. In addition, Low Boom analysis capability was added to the process. Engineering 

disciplines representing Advanced Design are Configuration, Aerodynamics, Stability & Control, Propulsion, Mass 

Properties, Acoustics, and Performance. 

Figure A.2.2 shows disciplinary modules and tools used in this process. Fidelity level of analysis is defined in 

methods used by each discipline. Disciplines shown in Figure A.2.3 are coordinated to ensure that the appropriate 

analysis match the intent of the study. 

ModelCenter is used to integrate and automate required discipline modules. ModelCenter is licensed Phoenix 

Integration software intended for a modern Multi-Disciplinary environment. The ModelCenter model represents the 

fusion of inputs and analysis codes used in aircraft performance analysis and sizing. Amount of integration is set by 

engineering Subject Matter Experts (SME). This environment requires engineers to interact with results and ensure 

correct automation. In this context, automation is relative and requires engineering judgment. 

Configuration is responsible for coordinating the flow of geometry information to engineering disciplines shown in 

Figure A.2.4. GeoDuck (Boeing parametric geometry program) was used to provide geometry inputs to the 

disciplines. Rules for generating geometry, inputs, and constraints are determined by SME using study 

requirements. Inputs on wing structure for Mass Properties, planform characteristics and airfoil definition for 

Aerodynamics are inputs needed by the analysis team. Another example of inputs is fuselage cross-sectional area 

constraints for minimum wave drag area-ruling by Aerodynamics as shown in Figure A.2.5. A brief description of 

GeoDuck is shown in Figure A.2.6. 

Parametric geometry models representing attributes for Supersonic NRA N+2 were created for analysis shown in 

Figure A.2.7. GeoDuck output IGES files in ModelCenter enable the creation of analysis input datasets. Creating 

UDP (vortex lattice code) input datasets are examples shown on Figures A.2.8 & 9. 

Aerodynamics is responsible for providing a starting linear design and Aerodynamic Characteristics for 

Performance. The aerodynamic module consists of low speed and high speed. Figure A.2.10 shows a description of 

the high speed module and associated tools for Aerodynamic buildups. Supersonic high speed drag buildups use 

Linear Theory plus nonlinear corrections from existing databases with provisions for technology projections. 

Aerodynamic codes wrapped in ModelCenter are TEA80 (wave drag), A389 (supersonic lift drag), and Dragj (high 

speed buildup). Example data release is shown as a Mach*Lift/Drag (ML/D) plot. The linear design uses camber 

definition obtained from A389 and area-ruled fuselage from TEA80. 

Low speed aerodynamic module description is shown in Figure A.2.11. Similar to high speed the buildups use 

Linear Theory corrected by applicable databases. UDP is the vortex lattice code wrapped for low speed. Released 

aerodynamic data are shown as lift curves and drag polars for different high lift device deflections. 

Stability and Control module is currently limited to tail volumes. Method activities have been initiated to 

incorporate static stability and time to double capabilities. 
214 

Propulsion module description is shown in Figure A.2.12. Initial capability allows engine decks to load 

automatically for Performance. Propulsion inputs are passed to the Propulsion module where scaling rules are 

invoked during Performance vehicle and engine sizing. Wrapping NPSS into the module under ModelCenter is a 

current BR&T method activity with preliminary implementation. 

Mass Properties is responsible for Weights release to Performance. QWIKO (Supersonic Weights Estimation) 

shown in Figure A.2.13 was wrapped in ModelCenter. Mass Properties inputs are generated by GeoDuck and 

passed to QWIKO for weights buildup. Weights statement is automatically generated with appropriate outputs 

linked to Performance. 

Performance module description is shown in Figure A.2.14. This module includes script wrappers to automate the 

flow of inputs required for Performance analysis. Takeoff and mission programs are operational with Landing in 

beta version. ModelCenter is setup for stand alone Performance run or Design of Experiments (DOE). DOE 

provides parametric results that could be reviewed in Data Explorer, exploring the design space with visualization 

tools. 

In ModelCenter links are setup to iterate on Mission takeoff gross weight (TOGW) with the Weights module 

providing an initial guess. Scaling rules have been setup in the discipline modules to allow wing area and engine 

thrust scaling. Takeoff Performance noise constraint in terms of jet exit velocity, rate of climb at the transonic pinch 

point are all captured for design sensitivity trades. These trades are captured through DOE. 

Low Boom configurations are enabling attributes for a viable supersonic aircraft. Supersonic NRA N+2 requires an 

exploration in design for Low Boom solutions. A Low Boom analysis module shown in Figure A.2.15 has been 

incorporated in the MDA ModelCenter model. This module does not have the same level of automation as the 

Performance modules. Difficulties in automation are due to the level of fidelity for Aerodynamic analysis. 

Addressing low boom design requires at least an Euler level solution for the initial pressure signature. 

The low boom module incorporates MDBOOM (Boeing propagation code) in ModelCenter. GeoDuck IGES 

surface geometry shown in Figure A.2.16 is passed to MADCAP and ICEM for surface preparation and mesh 

generation shown in Figure A.2.17. Scripts are used to automate CART3D submits to obtain a solution at the cruise 

lift coefficient. CART3D pressure contours at the cruise lift coefficient are shown in Figure A.2.18. Lift and area 

distribution of the CFD solution shown in Figure A.2.19 are used to calculate the F Function shown in Figure 

A.2.20. The F Function is the final far-field signature that MDBOOM propagates through the atmosphere. Figure 

A.2.21 shows the sonic boom waveform at the ground and perceived loudness (noise metric). 

Figure A.2.22 shows the summary and status of the MDA capability. The ModelCenter model has been 

demonstrated on the 765-072 Type Species with DOE. Sensitivity studies have been done at various level of 

automation. Low boom analysis has been exercised on converged Performance airplanes. 

Further work needs to be done to mature models for robustness. It is difficult to have robust parametric geometries 

while producing high quality surfaces for CFD. Further work on low boom shaping is continuing for higher order 

analysis. 
215 

Figure A.2.1 MDA Capability Outline. 

Figure A.2.2 MDA Tools Suite Components. 

• Multi-Discipline Analysis Architecture 

– Configuration 

– Aerodynamics 

– Propulsion 

– Mass Properties 

– Performance 

– Acoustics (Vjet=1100 fps) 

– Stability and Control (Volumes) 

• Low Boom Shaping 

– MDBOOM integration 

– Low boom design 

• Summary and Status 
216 

Figure A.2.3 Supersonic ModelCenter Model. 

Figure A.2.4 Geometry Model Description. 
217 

Figure A.2.5 Geometry Module Inputs. 
218 

Figure A.2.6 GeoDuck Environment. 

Figure A.2.7 Parametric Model Species. 
219 

Figure A.2.8 765-070 Type Species UDP Modeling of Parametric Model IGES output. 

Figure A.2.9 765-072B Type Species UDP Modeling of Parametric Model IGES output. 
220 

Figure A.2.10 High Speed Aerodynamic Module produces released data for Performance. 
221 

Figure A.2.11 Low Speed Aerodynamic Module produces released data for Performance. 
222 

Figure A.2.12 Propulsion Module Description. 

Figure A.2.13 Mass Properties Module Description. 

QWIKO is used for Mass Properties 

Mass Properties inputs are generated by the GEODUCK 

model and are passed to QWIKO for weights buildup 

Takeoff gross weight is cycled between QWIKO and 

mission codes until the two values converge 

NPSS: Numerical Propulsion System Simulation 

Given cycle parameters, NPSS produces an engine deck 

which can be read by performance codes. 

Engine/nacelle is scaled based on assumed trends. This 

information is passed to the GEODUCK geometry engine. 
223 

Figure A.2.14 Performance Module produce Takeoff and Mission results. 

Figure A.2.15 Low Boom Module Description. 

• GEODUCK passes design IGES files 

 to MADCAP and ICEM 

2. Grids generated for CART3D 

• Three angle of attacks submitted 

 to CART3D to get lift coefficient at cruise 

• Under-track area and lift distribution 

 obtained from CART3D solution 

• F-Function calculated and propagated 

 for ground signature and loudness metrics 

AD_process: Compiles and loads data from 

Aerodynamics 

AP_mission: Mission performance code (Model 

converges take-off gross weight with Mass 

Properties) 

AP_takeoff: Takeoff performance code 

Takeoff Printout 

Mission Printout 
224 

Figure A.2.16 Export IGES for Meshing. 

Figure A.2.17 Surface Mesh for CFD. 
225 

Figure A.2.18 CART3D (CFD) Solution for Sonic Boom Propagation. 

Figure A.2.19 Area and Lift Distribution for F Function analysis. 

0

100

200

300

400

500

600

0 25 50 75 100 125 150 175 200 225 250

X (ft)

AEQ (sqft)

AREA

LIFT

TOTAL

Cp contours 

Mach= 1.6 

h= 48,200 ft. 

Wt.= 262,000 lbs. 

CL= 0.154 
226 

Figure A.2.20 Generated F Functions for Sonic Boom Propagation. 

Figure A.2.21 Ground Signature and Loudness results. 

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

0.05 0.10 0.15 0.20 0.25 0.30 0.35 0.40 0.45

T (sec.)

P (psf)

SOURCE: MDBOOM 

PldB = 108.8 

-1.25

-1.00

-0.75

-0.50

-0.25

0.00

0.25

0.50

0.75

0 25 50 75 100 125 150 175 200 225 250 275 300

X (ft)

F(X)

AREA

LIFT

TOTAL

SOURCE: MDBOOM 
227 

Figure A.2.22 Summary and Status. 

 

• Multi-Discipline Analysis Architecture has been 

integrated in ModelCenter 

– Demonstrated in sensitivity studies 

– Beta version 

• MDBOOM integration 

– IGES to CART3D completed 

– Area and lift distribution in checkout 

– F-Function, signature and loudness metrics in checkout 

• Tasks to be completed 

– Model calibration 

– NPSS / WATE integration (option 2) 

– MDBOOM / Signature Shaping 
228 

Appendix 3 - GIT and R-R Non-Proprietary Technology Roadmapping 

The MFTF engine cycle selected by Boeing for the 076E configuration fails to meet the vehicle range requirement 

of 4000 nmi. In order to meet the range requirement Boeing determined that the vehicle fuel burn needed to be 

reduced by approximately four percent. Georgia Tech performed a technology gap analysis in order to determine 

what level of engine technology would be needed to meet mission goals. The process for the technology space 

exploration is shown in Figure A3.1. First the technology level is varied on the baseline cycle in order to determine 

the level of technology needed. After the technology level has been selected and implemented the engine cycle is reselected to find the best solution. 

Technology Space Exploration 

The first step is to identify the technology level needed to achieve the mission goals. This was done by taking the 

engine cycle selected for the 076E, from now on referred to as the baseline cycle, adjusting the design point 

efficiencies and analyzing the results. Since the technology space exploration was performed without a vehicle 

mission analysis, TSFC at a constant mid cruise thrust was used as a surrogate for fuel burn. Mid cruise was defined 

as Mach 1.6 at approximately 51,800 feet for the purposes of this exploration. The NPSS MFTF model was 

modified to allow for “technology factors” to be applied. These technology factors are scalars on the design point 

efficiencies of the turbomachinery components. Technology effects on the inlet and nozzle were not investigated. A 

design of experiments was run to determine coefficients bi and bij of the Response Surface Equations (RSE’s) for the 

technology space. RSE’s are a multivariate linear regression that assumes a Taylor series second-order 

approximation of the response, as follows: 

Equation 1. Response surface equation (RSE). 

A standard least squares regression is used to determine the coefficients in the RSE based on a given set of data 

generated intelligently through a Design of Experiments. 

A DoE was executed using the ranges of technology scalars used are shown in Table A3.1. The baseline cycle was 

held constant during the DoE execution; however, thrust was allowed to vary so that the engine could scale to meet 

vehicle climb and cruise thrust requirements as the thrust lapse rate changed. 

Once the DoE data was available RSE’s were constructed and used to create a prediction profiler, as shown in 

Figure A3.2. The profiler allows the designer to easily obtain sensitivities, see potential tradeoffs, and rapidly 

explore the design space that allows instantaneous results without having to re-execute the entire NPSS model. 

Several considerations went into selecting the resulting technology factors listed in Table A3.2. Technologies level increases for 

the various components were evenly distributed as much as possible. The purpose of this was to try to capture a gradual increase 

in all technology levels which is more a more feasible outcome than a drastic increase in one and no increase in the others. Also, 

since the HPC efficiency has such a large impact on the thrust lapse rate it was used not only to reduce SFC, but it’s value was 

somewhat limited by the requirement to meet climb thrust requirements. Finally, since the technology needs to be ready for IOC 

in 2020, engineers at Georgia Tech decided to limit the maximum increase in component efficiency to 2%. The baseline cycle 

TSFC reduction was reduced by 2.2% using the selected technology level in Table A3.2. This does not meet the 4% requirement; 

however, the effect of higher component efficiency on engine performance is not independent of the cycle. For this reason the 

cycle needed to be reselected to attain the maximum possible performance increase. 

Cycle Design Space Re-Exploration 

Once the appropriate technology levels were selected another DoE was run. This time a Latin Hypercube type of 

DoE was executed. It maximizes the distance between DoE points in the multidimensional design space. The ranges 

used for the cycle design space exploration are the same as those used to originally select the baseline cycle. The 

cycle which met thrust requirements and had the best TSFC improvement was selected from the DoE results. This 

cycle, which represents a solution in the discrete design space, was then optimized for cruise TSFC using a gradient 

based optimizer. The HPC efficiency and CCA temperature drop were also allowed to vary in this gradient 
229 

optimization since they have a large correlation with the cycle parameters on thrust lapse. The resulting cycle and 

performance is shown in Table A3.3. 

Non-Proprietary Technology Gap Analysis Conclusions 

By increasing the technology level and re-optimizing the cycle a 3.6% reduction in TSFC was achieved relative to 

the baseline 076E cycle. If this same process were to be performed with a parametric vehicle in the loop then it may 

be possible to achieve the 4000 nmi mission range requirement. Also worth noting is that the resulting efficiencies 

listed in Table A3.3 are not excessively higher than today’s levels. The HPC and HPT efficiencies are reasonable 

and since a multi-stage fan is assumed the low FPR means that two lower pressure ratio, more efficient, fan stages 

could be used. 

Figure A3.1 Process for Technology Selection 

Table A3.1 Technology Factor Ranges for Technology Space Exploration 

Inputs Range

Fan poly eff scalar 1.00 – 1.03 

HPC poly eff scalar 1.00 – 1.03 

HPT adia eff scalar 1.00 – 1.03 

LPT adia eff scalar 1.00 – 1.03 
230 

Figure A3.2 Technology Space Prediction Profiler 

Table A3.2 Selected Technology Factors 

Inputs Value Design Point Eff 

Fan poly eff scalar 1.02 0.923 (polytropic) 

HPC poly eff scalar 1.008 0.922 (polytropic) 

HPT adia eff scalar 1.015 0.934 (adiabatic) 

LPT adia eff scalar 1.01 0.939 (adiabatic) 
231 

Table A3.3 Results of Technology Gap Analysis 

Engine Cycle and Technology-Factor Settings 

 Baseline Technology Infused 

FPR 2.21 2.20 

OPR 39.21 37.60 

Extraction Ratio 0.9 0.99 

T41 Max (°R) 4,000 3,391 

Throttle Ratio 1.100 1.087 

SLS Net Thrust (lbf) 41,133 44,422 

k_effFan 1.000 1.020 

k_effHPC 1.000 1.015 

k_effHPT 1.000 1.015 

k_effLPT 1.000 1.010 

CCA ΔT (°R) 0 125 

Output from Engine Model 

Takeoff Thrust (lbf) (Mach 0.255 / 

1000 ft) 39.826 42,605 

Takeoff Exit Jet Velocity (ft/s) 1,516 1,448 

Fan Spec Flow (lbm/s) @ SLS 40.0 42.75 

Mid-Cruise TSFC lbm/(lbf*hr) 0.9611 0.9260 

Fan Polytropic Efficiency 0.9050 0.9230 

HPC Polytropic Efficiency 0.9150 0.9287 

HPT Adiabatic Effciency 0.9200 0.9338 

LPT Adiabatic Effciency 0.9300 0.9393 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Contractor Report

 4. TITLE AND SUBTITLE

N+2 Supersonic Concept Development and Systems Integration

5a. CONTRACT NUMBER

NNL08AA16B

 6. AUTHOR(S)

Welge, Harry R.; Bonet, John; Magee, Todd; Chen, Daniel; Hollowell, 

Steve; Kutzmann, Aaron; Mortlock, Alan; Stengle, Josh; Nelson, Chet; 

Adamson, Eric; Baughcum, Steve; Britt, Robert T.; Miller, Gregory; Tai, 

Jimmy 

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

NASA Langley Research Center The Boeing Company

Hampton, VA 23681-2199 5301 Bolsa Avenue

 Huntington Beach, CA 92647-2048

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

Langley Technical Monitor: Lori P. Ozoroski

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 05

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

Supersonic airplanes for two generations into the future (N+2, 2020-2025 EIS) were designed: the 100 passenger 765-072B, 

and the 30 passenger 765-076E. Both achieve a trans-Atlantic range of about 4000nm. The larger 765-072B meets fuel burn 

and emissions goals forecast for the 2025 time-frame, and the smaller 765-076E improves the boom and confidence in 

utilization that accompanies lower seat count. The boom level of both airplanes was reduced until balanced with performance. 

The final configuration product is two “realistic”, non-proprietary future airplane designs, described in sufficient detail for

subsequent multi-disciplinary design and optimization, with emphasis on the smaller 765-076E because of its lower boom 

characteristics. In addition IGES CAD files of the OML lofts of the two example configurations, a non-proprietary parametric 

engine model, and a first-cycle Finite Element Model are also provided for use in future multi-disciplinary analysis, 

optimization, and technology evaluation studies. 

15. SUBJECT TERMS

Advanced concepts; Conceptual Design; Supersonic; Systems Analysis

18. NUMBER

 OF 

 PAGES

232

19b. TELEPHONE NUMBER (Include area code)

(443) 757-5802

a. REPORT

U

c. THIS PAGE

U

b. ABSTRACT

U

17. LIMITATION OF 

 ABSTRACT

UU

Prescribed by ANSI Std. Z39.18

Standard Form 298 (Rev. 8-98)

3. DATES COVERED (From - To)

03-2008 to 08-2009

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

NNL08AA35T

5f. WORK UNIT NUMBER

984754.02.07.07.12.05

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/CR-2010-216842

16. SECURITY CLASSIFICATION OF:

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, 

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this 

collection of information, including suggestions for reducing this burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and 

Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person 

shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number.

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS.

1. REPORT DATE (DD-MM-YYYY)

01- 08 - 2010