NASA Contractor Report 4236

[

Study of High-Speed

Civil Transports

Summary

Douglas Aircraft Company

New Commercial Programs

Long Beach, California

Prepared for

Langley Research Center

under Contract NAS1-18378

N SA

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Division

1990
__ix!_._r ̧

_I
'l

CONTENTS

Summary ...................... ................................................................. xi

Introduction .................................................................................... 1

Market Analysis ................................................................................ 3

Vehicle Concepts ............................................................................... 9

Performance and Economics ...................................................................... 23

Airports and Environment ...................................................................... 35

Conclusions. .................................................................................. 45

Recommendations .............................................................................. 47

References .................................................................................... 49

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4

FIGURES

1. High-Speed Civil Transport Study ........................................................ 1

2. International Passenger Traffic for Major Regions ........................................... 3

3. Passenger Traffic Forecast ............................................................... 4

4. Passenger Distribution .................................................................. 5

5. Top Pacific Region Markets .............................................................. 5

6. Readjustment to Time Changes. : ........................................................ 6

7. Aircraft Productivity and Utilization ...................................................... 6

8. Effect of Fare Premium on Market Capture ............................................... 7

9. HSCT Fleet Requirements .............................................................. 8

10. HSCT Study Concepts ........... ....................................................... 10

11. Engine Efficiency ...................................................................... 11

12. Advanced Subsonic Transport ............................................................ 12

12 13. Mach 3.2 Baseline Configuration (D3.2-3A) ..................................... : .........

14. Mach 3.2 Increased Span Configuration (D3.2-4B) .. ....................................... 13

15. Mach 3.2 Lift/Drag ..................................................................... 14

16. P&W Mach 3.2 Variable Stream Control Engine ........................................... 14

17. Mach 3.2 Maximum Temperatures ....................................................... 16

18. Mach 5.0 Configuration ................................................................. 17

19. Mach 5.0 Lift/Drag ..................................................................... 18

20. GE Mach 5.0 Variable Cycle Hypersonic Jet Engine ........................................ 19

20 21. Mach 5.0 Maximum Temperatures .......................................................

22. Mission Profile ......................................................................... 23

23. Key Overland Market New York-Tokyo ................................................... 24

24. Mach 3.2 Concept Sizings for Partially Subsonic Range ..................................... 25

25. Weight Breakdown ..................................................................... 26

26. Fuel Efficiency ......................................................................... 26

27. Economic and Market Study Model ......................................... : ............ 27

28 28. Fuel Cost .............................................................................

29. Aircraft Operating Costs ................................................................ 28

30. Annual Revenue ....................................................................... 29

31. HSCT Economics at Zero Fare Premium ................................................. 30

32. Mach 3.2 Concept Production Breakdown ................................................. 31

31 33. HSCT Worth/Price Comparison .........................................................

32 34. Mach 3.2 Economic Viability ............................................................

35. Mach 3.2 HSCT Benefits to U.S. Economy ................................................ 33

36. Projection of Potential Weight Reduction ................................................. 34

37. Aircraft Characteristics ................................................................. 35

38. Gate Parking .......................................................................... 36

39. Estimated FAR Part 36 Noise Levels (EPNdB) ............................................ 38

40. Mach 3.2 Concept 100 EPNdB Noise Contours ............................................ 38

41. Sonic Boom Waveforms ................................................................. 39

42. Mach 3.2 Concept Sonic Boom Overpressures on the Ground ............................... 40

43. Mach 5.0 Concept Sonic Boom Overpressures on the Ground ............................... 40

44. Mach 3.2 Sonic Boom Minimization ...................................................... 41

45. Mach 3.2 Engine Main Burner Emissions Index ..................................... _ ...... 43

46. Mach 3.2 Annual Fuel Burn ............................................................. 43

47 47. Technology Plan .......................................................................

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GLOSSARY

ADP

ADS

AERA

AI

A1 MMC

Alt

ASM

ASNM

AST

ATA

ATC

ATR

CAB

CCTV

CFD

cg

CR

dB

dB(A)

DOC

EINOx

EPNdB

EPNL

F

FAA

FAD

FAR

FC

FEM

g

GE

GNP

GPS

Graphite/HFC, TZM

GW

HR

HSCT

HSPA

IATA

ICAO

IN.

Isp

IVP

L_e

lb

LCdn

LFC

LH2

LNG

M

MDBOOM

Advanced Ducted Prop

Automated Dependent Surveillance

Automatic En Route ATC

aluminum

aluminum metal matrix composites

altitude

available seat-mile

available seat-nautical mile

Advanced Supersonic Transport

Air Transport Association

air traffic control

air turboramjet

Civil Aeronautics Board

closed-circuit television

computational fluid dynamics

center of gravity

cruise

decibel

unit of A-weighted noise level

direct operating cost

NOx Emissions Index (lb NOx per 1,000 lb fuel burned)

Effective Perceived Noise Level in decibels

Effective Perceived Noise Level

fahrenheit

Federal Aviation Administration

Fuel Advisory Departures

Federal Airworthiness Regulations

first class

finite element model

acceleration of gravity

General Electric Aircraft Engines

gross national product

Global Positioning System

Graphite Fibers with Hafnium Carbide Matrix Composites, Molybdenum

Titanium Zirconium alloy

gross weight

hour

High-Speed Civil Transport

high-speed propulsion assessment

International Air Transport Association

International Civil Aviation Organization

inches

specific impulse

inverted velocity profile

C-Weighted Sound Exposure Level in decibels

pounds

Day-Night Average C-Weighted Sound Exposure Level

laminar-flow control

liquid hydrogen

liquid methane gas

Mach number

Sonic Boom Analysis, Douglas Aircraft

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vii

++m, Ir r Ptrtom at:tY
GLOSSARY

(Continued)

MIT

MMLI

MMC

MTOGW

NAS Plan

NASA

NASP

NMI

NOx

OEW

P&W

PC

PEEK

PLdB

PM/PV

PMR

psf

R&T

RB/QQ

RPM

RSR

Sw

SCAR

SCR

SCS

SERN

SFC

SPF/DB

ST

T

TAD

TD Ni Cr

Ti

Ti MMC

TOGW

TPS

TSJF

T/W

USSR

VABI

VCE

VCHJ

VSCE

T

Ap

Massachusetts Institute of Technology

modularized multilayer insulation

metal matrix composite

maximum takeoff gross weight

National Airspace System Plan

National Aeronautics and Space Administration

National AeroSpace Plane

nautical mile

oxides of nitrogen (all species)

operator's empty weight

Pratt & Whitney

polymetric composites

poly ether-ether ketone

Steven Mark VII Perceived Level Of Loudness in decibels

premixed, prevaporized

polyimide resin

lb/ft 2

research and technology

rich burn-quick quench

revenue passenger mile

rapidly solidification rate

wing area

Supersonic Cruise Aircraft Research

Supersonic Cruise Research

silicon carbide system

single expansion ramp nozzle

specific fuel consumption

super-plastically formed, diffusion bonded

statute mile

temperature

technology availability date

thoria dispersion nickel chrome

titanium

titanium metal matrix composite

takeoff gross weight

thermal protection system

thermally stable jet fuel

(engine) thrust/weight ratio

Union of Soviet Socialist Republics

variable area bypass injector

variable cycle engine

variable cycle hypersonic jet

variable stream control engine

slenderness factor = volume/(wing area) 3/2

overpressure

viii

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FOREWORD

The High-Speed Civil Transport (HSCT) study integrated results of technical and economic analysis of

various aircraft to determine their commercial potential and corresponding technology requirements. This

extended beyond previous primarily technology-oriented activities such as the Advanced Supersonic Transport

(AST) and Supersonic Cruise Research (SCR), and included consideration of ongoing technology developments

of the National AeroSpace Plane (NASP) program. Appropriate technologies were assessed in terms of the

commercial value of HSCT aircraft.

Work was accomplished by Douglas Aircraft Company in Long Beach, California. This work commenced

in October 1986 at the direction of the NASA Langley Research Center, Hampton, Virginia, and was jointly

funded under Contract NAS1-18378.

The NASA Contracting Officer's Technical Representative was Mr. Robert W. Koenig during the initial

phases and Mr. Charles E. K. Morris, Jr., during the final phases; the Douglas program manager was Mr. Donald

A. Graf. Principal investigators were Mr. H. Robert Welge, aerodynamics, acoustics, and assistant program

manager; Mr. Gordon L. Hamilton, propulsion, fuels, emissions, and thermal; Mr. M. A. "Pete" Price, structures and materials; Mr. Bruce W. Kimoto, systems and weights; Mr. Richard T. Cathers, configuration integration; Mr. John W. Stroup, market research for the initial phases; Mr. Munir Metwally, economics and market

research for the final phases; Mr. Marc L. Schoen, airports; and Mr. Maurice Platt, manufacturing and development costing. McDonnell Aircraft Company technical staff provided consultation services on NASP technologies.

Other Douglas HSCT team members are:

Administration

Aerodynamics and Acoustics

Business Operations

Configuration

Economics and Market Research

Human Factors

Laminar Flow

Product Support

Propulsion

Structures and Materials

Weights

E. C. Anderson

G. A. Intemann, J. A. Harbauer, Dr. D. L. Antani,

R. A. Sohn, A. Mooiweer, A. K. Anderson,

G. S. Page, Dr. A. G. Powell, A. K. Mortlock

M. L. Shell

R. E. Lindenmuth

M. J. Berger

Dr. A. N. DeGaston

W. E. Pearce, N. M. Jerstad

R. E. Swartzbacker

F. R. Mastroly, J. A. Meyer, Dr. Y. Oida, C. R. Ross

Dr. E. G. Chow, D. E. Kerans, C. Y. Kam

G. J. Espil

z

C

ix

• SUMMARY

A systems study to identify the economic potential for a high-speed commercial transport has considered

technology, market characteristics, airport infrastructure, and environmental issues. Market forecasts indicate

a need for HSCT service in the 2000/2010 time frame conditioned on economic viability and environmental

acceptability. Design requirements focused on a 300-passenger, three-class service, and 6,500-nautical-mile

range based on the accelerated growth of the Pacific region. Compatibility with existing airports was an assumed

requirement. Mach numbers between 2 and 25 were examined in conjunction with the appropriate propulsion

systems, fuels, structural materials, and thermal management systems. Aircraft productivity was a key parameter with aircraft worth, in comparison to aircraft price, being the airline-oriented figure of merit. Aircraft screening led to determination that Mach 3.2 (TSJF) would have superior characteristics to Mach 5.0 (LNG) and the

recommendation that the next generation high-speed commercial transport aircraft use a kerosene fuel. q'hc

sensitivity of aircraft performance and economics to environmental constraints (e.g., sonic boom, engine emissions, and airport/community noise) was identified together with key technologies. In all, current technology

is not adequate to produce viable HSCFs for the world marketplace. Technology advancements must be accomplished to meet environmental requirements (these requirements are as yet undetermined for sonic boom and

engine emissions). High priority is assigned to aircraft gross weight reduction which benefits both economics

and environmental aspects. Specific technology requirements have been identified, which was the primc

objective of this study. National economic benefits are projected.

')

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xi
k
INTRODUCTION

?ii

Commercial aviation has grown to become a mature, highly competitive, worldwide industry. The growth

in air travel is projected to continue well into the 21st century with increasing demand for more efficient aircraft

and increasing focus on environmental compatibility. Manufacture of commercial aircraft is a significant element of the U.S. economy and a major contributor to the balance of trade. In 1985, the White House Office

of Science and Technology fostered long-range goals for continued U.S. leadership in aeronautics. In 1986,

NASA commissioned this 2-year High-Speed Civil Transport (HSCT) study of supersonic transport viability and

technology requirements to meet these goals.

This report is a summary of that work conducted by the Douglas Aircraft Company under contract

NAS1-18378. This study was initially of broad perspective and then focused on the critical aspects of the best

HSCT concepts with the overall goal of no measurable impact on the environment. At the outset, cruise speeds

ranging from Mach 2 to 25 were considered. Major elements included (1) vehicle technology assessment based

on all the traditional disciplines and their integration into vehicle concepts, (2) evaluations of HSCT operating

and production costs, and comparisons of vehicle price and worth, and (3) definition and assessment of the more

promising concepts in more detail. In addition, special factors such as airport infrastructure, fuel technology,

and the environmental concerns for sonic boom, community noise, and engine emissions were included.

Engine data were obtained through subcontracts to Aerojet TechSystems, General Electric Aircraft

Engines, and Pratt & Whitney (P&W). In addition, subcontracts have provided expertise in such areas as air

traffic control, cryogenic fuels technology, airport facilities development, energy analysis, sonic-boom technology, and market-related issues (such as schedules and passenger value-of-time). Airline viewpoints have been

incorporated from both U.S. and overseas airlines on an ad hoc basis.

The study was organized into three phases, as shown in Figure 1 with Phase I constituting a broad, goaloriented screening. At its conclusion, concepts utilizing hydrogen fuel were eliminated due to their relative low

economic performance caused by hydrogen fuel prices. Phase II focused on Mach 2.2, Mach 3.2, and Mach 5.0

without environmental restrictions for supersonic cruise over land, community noise, or accountability for

engine emissions/atmospheric interactions; this established baselines for follow-on evaluations. Concepts with

cruise speeds of Mach 3.2 and Mach 5.0 were selected for more detailed study in Phase III with the Mach 3.2

concept utilizing thermally stable jet fuel (TSJF) and the Mach 5.0 liquid methane (LNG). The focus of

Phase III was environmental acceptability. Since research may be based on the results of this work, it was considered imperative to assess Mach 3.2 and Mach 5.0 reflecting significantly different speeds and cruise altitude.

Phase III included integration of NASP technologies to access contributions from this ongoing program to civil

aviation.

NASA CONTRACT NAS1-18378

OCT 1986 AUG 1987 FEB 1988

1[ PHASE I I PHASE II I

I= 41 l= I[

10 MONTHS 6 MONTHS

• COMMERCIAL • SELECTED *

VALUE STUDY CONCEPTS

TRADE

• TECHNOLOGY STUDIES *

ASSESSMENT

AN'D VEHICLE

CONCEPT DEFINITION

• ENVIRONMENTAL

ASSESSMENT

PHASE III

8 MONTHS

ENVIRONMENTAL

FOCUS

TECHNICAL ASSESSMENT

UPDATE

• CONTINUED • COMMERCIAL VALUE

COMMERCIAL UPDATE

VALUE STUDY

ENVIRONMENTAL

DATA DEVELOPMENT

FIGURE 1. HIGH-SPEED CIVIL TRANSPORT STUDY

OCT 1988

I OPTIONS

• RELATED

TECHNOLOGY

STUDIES
Environmental acceptability and economic viability are key to technology requirements which will significantly influence U.S. industry's eventual decision to go ahead on an HSCT development and production program. Douglas considers resolution to the environmental issues to be of first-order importance with selection

of a specific cruise Mach number a part of the HSCT design process.
MARKET ANALYSIS

Assessment of the commercial value of the HSCT requires a comparison between its economic worth and

cost-based price. This includes the following:

• Traffic demand • Cash flow

• Market share • Airline return on investment

• Passenger value of time • Tax law

• Productivity • Useful life

• Aircraft scheduling • Quantity

• Fleet sizes • Development cost

• Passenger revenue • Recurring production cost

• Operating costs • Manufacturers' return on investment

The overall evaluation process, specially designed economic and operational modes, and supporting data

were the products of this study. Frequent consultation with airlines, including Northwest, ,Federal Express,

American, Delta, United, Pan Am, Japan Air Lines, Alitalia, and British Airways, provided invaluable

exchanges regarding traffic projections, schedules, economic parameters, and related matters. Expert consultation on such issues as aircraft scheduling, utilization, and productivity as well as passenger value of time was

received from Massachusetts Institute of Technology, Purdue University, and Quinnipiac College.

Douglas-developed econometric methods were used to forecast traffic through the year 2000. The traffic

estimates for the 18 international IATA regions were the data source. Further considerations of range, traffic,

and mileage over land reduced the base to 10 regions. These 10 regions were used as the arena of competition

between the HSCT and an advanced subsonic transport reference vehicle. These regions are primarily intercontinental, over water, and long range. Four regions comprise between 85 to 90 percent of the total international traffic (Figure 2). The North Atlantic and North and Mid-Pacific are the major traffic regions. In 1986,

the North Atlantic recorded more than twice the revenue passenger miles compared to the North and MidPacific. By the year 2000, equal traffic is expected in these two major sectors. Overall, traffic is predicted to total

446 billion available seat-miles (ASM), with 53 percent being over water routes. These projections are based

on individual country economics, trade characteristics, and priorities.

EUROPE/FAR EAST

AUSTRALASIA

NORTH AND

MID-PACIFIC

NORTH

ATLANTIC

2000]

1986]

INTRA

FAR EAST

AUSTRALIA

_4 13_

EUROPE/ INTRA NORTH NORTH

FAR EAST FAR EAST MID-PACIFIC ATLANTIC

RPMs IN BILLIONS

FIGURE 2. INTERNATIONAL PASSENGER TRAFFIC FOR MAJOR REGIONS
For the period of 2000 through 2025, three economic scenarios were considered: (1) uninhibited growth

at the rates in the final years of the 1986-2000 time period, (2) continued rapid growth in the Pacific Rim countries with rates moderating in the latter half of the 25-year period, and (3) growth rates decreasing to the levels

of the projected rates in the general economies of the regions. Of the three scenarios considered, the second

with a growth to 2,386 billion seat-miles is considered most likely (Figure 3). This represents five times the traffic

projected for the year 2000. These data represent a first-order market assessment and basis for productivity and

utilization analysis; and they do not represent an exhaustive research for HSCT opportunities.

ANNUAL

SEAT-MILES

(BILLIONS)

4,000

3.000

2.000

1.000

"GREATER THAN

2,000 STATUTE MILES 3,521

446

]

1986 2000

2,386

]

/ 1,852

2025

YEAR

FIGURE 3. PASSENGER TRAFFIC FORECAST

Based on this market analysis and previous Douglas supersonic aircraft studies, a vehicle size of approximately 300 seats is necessary for competitive economics; therefore, all vehicles were sized accordingly. The passenger market is segmented into four fare classes; first class, business, and coach including full fare and discount

fare base. First class, business, and approximately half of the full-fare coach market travel primarily for business

reasons and the remainder travel for personal or pleasure reasons (Figure 4). It is assumed that the HSCT will

be operated based on current airline business practices of fare management and thus accommodate all levels

of passenger traffic. Correspondingly, HSCT economic analyses include fare distinction by class of service.

Range requirements will be mandated by the key Pacific markets with an eye to the new long-range subsonic transports now emerging. Ranges vary from slightly less than 4,000 statute miles for the Honolulu-to-Tokyo market, to approximately 5,500 statute miles for the Los Angeles-to-Tokyo market, to nearly 7,000 statute

miles for the New York-to-Tokyo market (Figure 5). Los Angeles to Sydney, with 7,500 statute miles, represents

the upper nonstop range requirement. This range capability captures 80 percent of all long-range, nonstop traffic, and thus was adopted as a design requirement. In addition, other design objectives were established based

on the assumption that for an HSCT to be viable, it must be operable from current airports. This included a

maximum takeoff field length of 11,000 feet and maximum takeoff gross weight of 1,000,000 pounds. Approach
ASMs

(BILLIONS

PER

YEAR)

140

120

100

80

6O

40

20

3

BUSINESS

50%

YEAR 2000

(30./W

\ / F_)_CH\ /

PERSONAL

50%

FIGURE 4. PASSENGER DISTRIBUTION

LAX-N RT

P

HNL-NRT F

v--r-f I I

4 5

LAX-SYD

HKG-SFO .j

JFK-NRT _-J

I I

6 7

DISTANCE (1,000 ST MI)

FIGURE 5. TOP PACIFIC REGION MARKETS

speeds of 140 knots will maintain compatibility with ATC. Accelerations not to exceed 1/10 g during climb will

assure passenger comfort. Environmental acceptability is a key requirement.

Market stimulation by HSCT flight-time savings is considered to be a significant advantage by airlines.

There is positive evidence that the advent of jet travel and the resulting time savings caused accelerated market

growth particularly over the long-distance routes. Airlines consulted during this study believe a viable HSCT

could cause similar market response but no attempt was made to quantify levels. The favorable affects of HSCT

flight-time savings when crossing many time zones will reduce jet lag. Jet lag encompasses the feelings of irritability, mental and physical lethargy, disorientation, and fatigue. Adjustment for the jet lag occurs in about the

time that it takes for the sleep-dependent and physical parameters to fall back into phase. The sooner a passenger arrives in the new time zone, the sooner the readjustment begins; a matter of hours is significant. Figure 6

shows the significant reduction to time changes achievable through flight-time reductions. Round trips in one

day should not create jet lag problems for those business travelers who take advantage of the short travel times

nor for HSCT crews.

5

::::
RESIDUAL

JET LAG

(HR)

10

HOURS _

\

68--_ .....

" " " " " " _ TIME ZONE CHANGES AND/OR

-- " - . _ _ ARRIVAL TIME DIFFERENCES

2 _..:::.....:::.....:::.....:::.....:::.....:::.....:::.....:::...._::.....:_...:_...._...._:...,.:..,::._:::.._....:_.::......::.........

......'"' A,,C,_,E pTA,B.,L,E............................... '"' '"' '"' '"' '"" '"" '" '""' _' '" ""'_'""'""_'" ' ,' ..... ' '" _,"_.7,,"_",,, ,,'"_,,'",,'",,_,,'" ,,'",,",,,,,,'" .......... ,,'",,'",,'",,'",,'",,,,,,'" ..................... ,,'",,'"/,,,'" ,,,'",/,,,'" ,,,'",,,'",,,'",.....,,

....'"''"' '"''l"'""'""'""4 ,'"','"'""'"' '" "' '"' '"'"" '""'""'""'"' '"''"' '""" " ,", ...... ....... _' ,"_" '" '"_''_ ,""','""'"' '""'""'""' o ..... ,,,,,,, ,,,,,.,,,, ,,,,, ,,,,, ,,,,,,,,,,,,,,,,,,I,,,,,,,,,,i,,,,,,,,,,,,,,t,,,,,,,,,,,.,f...................... _ ............... , , ,..... ,

1 2 3 4 5 6 7 8 9 10 11 12

DAYS AFTER ARRIVAL

FIGURE 6. READJUSTMENT TO TIME CHANGES

Market research highlighted more than 300 international city-pairs as candidates for HSCT service. A

global concept airline routing system consistent with current airline scheduling for large transport aircraft was

developed to determine aircraft use considering HSCT speed potential, real-world constraints such as airport

curfews, and passenger preferred times of day for travel. From this, aircraft productivity (seat-miles-per-year)

was determined based on cruise Mach number (Figure 7) assuming a 2-hour aircraft turnaround time and market applications ranging from 2,000 to 6,500 nautical miles.

PRODUCTIVITY

(SEAT-MILES/YEAR)

BILLION)

2.50

2.00 --

1.50 --

1.00 --

0.50 --

m

I

PRODUCTIVITY

I I I I I

1 2 3 4 5

CRUISE MACH

14

12

10

8

UTILIZATION

6 (HOURS/DAY)

4

2

0

FIGURE 7. AIRCRAFT PRODUCTIVITY AND UTILIZATION
Scheduling results showed the expected increase in productivity as cruise speed increased. In the Mach 4

or 5 range, productivity gains begin to diminish; and above Mach 6, very little additional productivity is achievable. Thus, Mach 5 to Mach 6 represents the commercial upper limit within the current air transportation system of nonstop, point-to-point service. Figure 7 includes the resulting aircraft daily utilization as a function of

cruise Mach number.

Passenger value of time, fare premiums, and blocktime differences between the HSCT and competing

subsonic aircraft affect market share. The underlying assumption is that a traveler will be an HSCT passenger

if the monetary value of the time savings exceeds the fare premium. The passenger market is segmented into

four different fare classes ranging from first-class travelers with a relatively high value of time to a highly pricesensitive aft cabin or discount coach market.

Figure 8 shows the potential traffic assuming no supersonic cruise over land will be allowed and the significant effect of fare premium on market capture. The data reflect subsonic cruise range capability.

°°l 400.000

ASMS

(MILLIONS)

YEAR 2000

NO SUPERSONICCRUISE

OVERLAND

300,000

200,000

100.000

/-- DISCOUNT COACH

FIRST-_

I I I I

0 10 20 3O 40

FAREPREMIUM

(PERCENT)

FIGURE 8. EFFECT OF FARE PREMIUM ON MARKET CAPTURE

50

World passenger jet aircraft requirements (excluding the USSR) are expected to total about 10,000 aircraft

in 2000 with about 1,800 aircraft in the long-range (greater than 3,500-nautical-mile range) class. Approximately

one-half of this long-range market is potential for the HSCT system. The HSCT with no fare premium may take

the place of a maximum of 900 aircraft and since the HSCT is some 2.5 to 3 times as productive as a subsonic

aircraft of the same size; 300 to 450 HSCT could do the work of 900 subsonic aircraft. If high-speed cruise over

land is prohibited, then requirements will be reduced by as much as 20 percent. If premium fares are required,

then the fleet sizes are further reduced. The HSCT fleet requirements in the year 2000 are shown in Figure 9

for the eight international (IATA) region system considered appropriate in view of the supersonic cruise over

land limitation. The share of the long-range market depends on fare premium and, ultimately, on the operating

costs.
NUMBER

OF

AIRCRAFT

800

600

40O

20O

ADVANCED SUBSONIC s "_ ""

TRANSPORT FLEET.,. _"

1t/'"/_ _ 1_/ YEAR 2000

f

/ LOST MARKET

"" "" _" _ MACH 3.2 FLEET

I I I I

0 10 20 30 40

FARE PREMIUM (PERCENT)

FIGURE 9. HSCT FLEET REQUIREMENTS

5O
VEHICLE CONCEPTS

The initial study efforts consisted of a broad screening analysis of the Mach 2 to 25 range. The development

of the HSCT benefitted from related work. This includes the supersonic transport work of the 1960s; the AST

activity of the 1970s, follow-on research by NASA, the Concorde, and the NASP. HSCT prime design objectives

are: accommodation of 300 passengers and a design range of 6,500 nautical miles (7,500 statute miles). In addition to passenger baggage, 500 cubic feet of volume is provided for 5,000 pounds of cargo. As a U.S. civil aircraft,

the HSCT will be certified and operated per FAA code of Federal Regulations. The design and certification will

conform to Part 25, and it will be operated under Part 91 and Part 121.

During the early study phases, a large number of candidate engines Mach number/fuel combinations were

evaluated. This included both military and commercial concepts and ranged from preliminary designs developed

for the high-speed propulsion (HSPA) studies for the Air Force, to commercial engines specifically tailored to

the Douglas high-speed civil transport configurations. Engine data have been supplied by Aerojet TechSystems,

General Electric Aircraft Engines, and Pratt & Whitney. All studies assumed an aircraft certification date of

21300/2010, with a corresponding technology availability date (TAD) of 1995/2000.

For high-speed applications, the choice of fuel is of broadened importance: the fuel energy content influences the size and weight of the airplane; the heat sink capability and thermal stability limits of the fuel influence

the Mach number. Some of today's fuel supply has been shown to have thermal stability limits above that of

Jet A (today's commercial standard fuel). This indicates that kerosene-based fuels can be produced with

enhanced thermal capability. The cost of the fuel becomes a more predominant factor in the operating

economics.

HSCT fuel system design is similar to that of the conventional subsonic airplane for cruise speeds up to

Mach 2.0 or 2.5. At higher cruise speeds, the requirements for an enhanced thermal stability kerosene fuel or

for LNG fuel may require extensive changes in the fuel system configuration. Aircraft gas turbine engines have

been designed and operated routinely with kerosene fuels in both commercial and military airplane service.

They also have accumulated millions of hours of operation on natural gas fuel in dual fuel marine and industrial

applications. Commercial jet engines experience fuel temperatures on the order of 325°E The SR-71, using

JP-7 fuel, is able to accommodate fuel temperatures as high as 600°E Figure 10 presents a matrix of vehicles

and fuels considered.

An advanced subsonic transport incorporating technologies expected in the year 2000 and using conventional Jet A fuel was defined to provide an economics baseline. Mach 2.2 and Mach 4 vehicles used a kerosenetype fuel with higher thermal capability. In addition, the Mach 4 studies also included endothermic and cryogenic fuels. Endotherrnic fuels were eliminated because of price and availability. Liquid methane offers a

16-percent higher energy per pound; however, it has an energy density only 60 percent of that of kerosene-type

fuels. Liquid hydrogen (LH2) also offers an increase of energy (180 percent) at a still further reduced energy

density (25 percent). Initial concept evaluations included comparisons of aircraft worth (to the operator) and

flyaway price. The hydrogen-fueled concepts fell short in aircraft worth because of high LH2 fuel cost and thus

were dropped from further consideration.

The following Mach numbers were selected for more detailed study with the rationale for their selection

as follows:

• Mach 3.2

-- Higher productivity than Mach 2.2 -- Aggressive technologies

-- Upper limit of kerosene-based fuel

• Mach 5.0

-- Highest productivity -- More aggressive technologies including NASP

-- Application of methane fuel

Commercial airplane operations beginning in the 2000/2010 time period focus these studies to kerosene:

based jet fuels and LNG. It is generally agreed that the thermal stability of Jet A is a limiting factor in high Mach

number applications. Based on inputs from refiners and engine manufacturers, JP-7 was chosen by Douglas as

the reference fuel for the Mach 3.2 studies. The properties of JP-7 that are desirable from the standpoint of

commercial application are related to thermal stability; however, some of the properties of JP-7 may not be

required. The term thermally stable jet fuel (TSJF) was adopted for this study.
MACH

C

= 0.85

MACH = 2.2

MACH = 6

VEHICLE MATRIX

JET A/TSJF

ENDOTHERMIC

LIQUID

FUEL NATURAL

TYPE GAS

LIQUID HYDROGEN

-0 0 0

- 0

- 0 0

- 0 0

I I I I

0.85 2.2 4 6

MACH NUMBER

O

I

12/20

FIGURE 10. HSCT STUDY CONCEPTS

One important parameter for preliminary screening was engine overall cruise efficiency. Overall efficiency is approximately equal to the product of the specific impulse times the flight velocity divided by the fuel

heating value. It is a measure of fuel energy conversion to jet kinetic energy and is the product of the thermal

efficiency and propulsive efficiency. Specific impulse, Isp (Isp = 3,600/SFC) is used in lieu of specific fuel consumption, SFC. Depending upon the design mission, fuel consumption during climb can amount to 25 percent

or more of the total fuel weight at takeoff. In the event that sonic boom restrictions preclude supersonic flight

over land, the aircraft may have to fly up to 40 percent of its design range at subsonic speeds. Thus the cycles

with the best supersonic cruise overall efficiencies, generally turbojets, will not necessarily be the optimum for

the vehicle. Engine cycle selection was made on the basis of vehicle takeoff gross weight and FAR Part 36,

Stage 3 takeoff noise requirements. The resulting engine efficiency levels are shown in Figure 11 compared to

a potential upper level. This evolved from studies of engine component performance and represents a variety

of engines and fuels. Engines developed for this study offer 20- to 25-percent improvement when compared to

the Concorde; however, exhibit efficiency levels significantly below the potential upper bound. It is expected

that more focused efforts will result in higher efficiency levels and corresponding improvement to vehicle

performance.

Advanced Subsonic Transport

Throughout the study, an advanced subsonic transport, incorporating projected technology improvements

for the year 2000, has been used for economic comparisons. The vehicle features an advanced wing, very-highbypass ducted fan engines burning conventional Jet A fuel and advanced systems. The configuration is shown

in Figure 12. The Mach 0.85 concept was based on Douglas Aircraft Company's development work for the

MD-11 and employs a shortened fuselage for consistency with a 300-passenger interior and an all new, highaspect-ratio supercritical wing with an aspect ratio of 11.41, and a span of 174.9 feet. Riblets were applied to

the wing, taft, fuselage, and nacelles. A hybrid laminar flow control system was included together with a wet

tail for cg control. The trimmed lift-to-drag ratio for the concept was 21.5.

10
OVERALL

CRUISE

PROPULSIVE

EFFICIENCY

(PERCENT)

7O

6O

5O

4O

3O

-- _ POTENTIAL

/ co co o 

2o I I I I I

0 1 2 3 4 5 6

CRUISE MACH NUMBER

FIGURE 11. ENGINE EFFICIENCY

The Mach 0.85 configuration incorporated application of newer materials and fabrication techniques.

SCS-8/AI 6061 aluminum metal matrix being the most efficient material and having both a high specific moduli

and specific compressive yield strength was used for the structure. To further reduce the weight, honeycomb

construction was used for the load-carrying skins instead of traditional skin-stringer design. It was estimated

that honeycomb construction would result in an approximate 15-percent weight reduction.

The P&W Advanced Ducted Prop (ADP) engine has been selected as a representative advanced subsonic

engine for long-range aircraft. The ADP is a two-spool, geared high-bypass, high-pressure ratio, ducted prop

engine with separate core (primary) and prop (duct or bypass) exhaust streams. One of its significant features

is a variable pitch ducted prop that provides good operability as well as reverse thrust. The cycle and component

design parameters for this engine are:

Overall pressure ratio 36

Fan pressure ratio 1.3

Bypass ratio 15

Maximum combustor exit temperature 2,650°F

Mach 3.2. The Mach 3.2 baseline concept, D3.2-3A (Figure 13), was based on the design work conducted

for the 1979 AST (Reference 1), under joint NASA/McDonnell Douglas funding. The configuration features

a double-sweep arrow planform wing, conical-taper fuselage, aft vertical and horizontal empennage, four P&W

duct-burning turbofan engines, TSJF, and a tricycle landing gear.

The fuselage was designed to accommodate a nominal seating arrangement of three classes: 10-, 30-, and

60-percent for first, business, and coach classes, respectively. The concept developed was unaffected by

constraints for sonic boom optimization. Supersonic drag necessitates use of a varying cabin cross section. The

fuselage incorporates single-lobe shaped cross sections; the maximum section is determined by a twin aisle with

11
'L : '

FIRST CLASS - 6 _ROSS AT 40 IN PITCH - 24 SEATS

COACH CI ASS - 9 ACROSS A1 34133 IN PITCH - 211 SEAIS

TOTAL - 301 _ATS

TOGW = 397.0(;0 POUNDS

OEW = 189,0(;OPOUNDS

5 W = 2680 FT 2

174 g _t"

184 | FT

/

628FT

I

FIGURE 12. ADVANCED SUBSONIC TRANSPORT

121.2 FT =

TOGW = 769.000 POUNDS

OEW = 231,335 POUNDS

SW = 9,500 FI 2

PAYtOAD

FmSlCtASS = 10f'[_(:ENI _ 4aCf_OSS^1.2iN PLICH _ JOSEAIS

aUStNESSCLASS _OV(IICEt_I = 46ACHOSSAt_tN Pa;C_¢ _OS(AIS

COACt¢CtASS _ 60_'[_(:eNl = '_AC_OSSA;321N P,ICH • le0S(AtS

IOtAt _ ImP|f_CtNI _ _OOSlAIS

FIGURE 13. MACH 3.2 BASELINE CONFIGURATION (D3.2-3A)

12 ORIGINAL PAGE IS

OF Poor QUALITY
seven-across coach seats, six-across first/business-class seats, and the minimum section is determined by a single

aisle, five-across coach seating arrangement or four-across business-class seats. All seat sizes are consistent with

those on MD-80 and MD-11 aircraft. The baseline interior arrangement provides service for 300 passengers

based on a maximum flight duration of 5 hours. Each class has its own galley, lavatories, coatrooms, and cabin

attendant stations. Four cabin doors per side are evenly distributed for rapid evacuation. Cabin windows are

incorporated.

The basic arrow-wing AST planform was modified with increased leading- and trailing-edge sweep to

improve supersonic performance at Mach 3.2 cruise. The wing design had a planform reference area of 9,500

square feet, an aspect ratio of 1.55, an inboard leading-edge sweep of 76 degrees and an outer panel leading

edge sweep of 62 degrees. The wing was optimized for a maximum wing-body trimmed lift-to-drag ratio. The

fuselage area distribution and camber were optimized for minimum wave drag at cruise conditions. The high-lift

system consisted of plain, trailing-edge flaps and full-span, simple-drooped leading-edge flaps. Laminar-flow

control (LFC) inboard of the wing planform break was limited by the fuel tank boundaries. Outboard of the

planforfia break, LFC extended to the flap-hinge line.

Longitudinal control and trim capability were provided by a totally movable horizontal surface with a

geared elevator. Ailerons and spoiler panels provided lateral control; directional control was provided by a

rudder.

A major environmental issue lies in the sonic boom resulting from supersonic cruise flight. A derivative

of the baseline configuration was developed to increase performance efficiency during subsonic flight over land.

This configuration, D3.2-4B, incomorates a wing planform modified to increase the span and reduce the leadingedge sweep of the outer wing panels (Figure 14). The maximum trimmed cruise lift-to-drag ratio of the Mach 3.2

configurations are shown in Figurc 15 for the subsonic and supersonic flight profiles over land. The difference

in the subsonic lift-to-drag ratios (shown for the -3A) was due to flight at lower altitudes.

TOGW = 880.000 POUNDS

DEW = 250,0C_ POUNDS

S w = 9.5_ FT 2

144 7F[

ln_l mien

31SF! ...................

•• ?

FIGURE 14. MACH 3.2 INCREASED SPAN CONFIGURATION (D3.2-4B)

The Mach 3.2 baseline engine is the P&W variable stream control engine (VSCE) duct-burning turbofan

using TSJE The VSCE (Figure 16) is an advanced, moderate-bypass-ratio, nonmixed-flow turbofan with duct

burner augmentation and a coannular nozzle with inverted velocity profile for jet noise reduction. P&W provided a preliminary assessment of takeoff noise reduction which identified the VSCE and the turbine bypass

13

ORIGINAL PAGE IS

OF POO_ QUALiTy
16

LIFT/DRAG

TRIMMED

12

._-3A

-- I_.,_.,.l _ ""lllll_

SUPERSONICOVERLAND FLIGHT PROFILE

' ,_ / -48 SUBSONIC OVERLAND FLIGHT FROFILE

-3A SUBSONIC OVERLAND FLIGHT PROFILE

I I I I I !

.5 1 1.5 2 2.5 3

MACH NUMBER

FIGURE 15. MACH 3.2 LIFT/DRAG

3.5

FIGURE 16. P&W MACH 3.2 VARIABLE STREAM CONTROL ENGINE

engine as having the greatest potential for satisfying FAR Part 36, Stage 3 requirements, with the VSCE having

the advantage due to its inverted velocity profile (IVP). A variable area suppressor allowed optimization of jet

noise at the sideline condition. In addition, an acoustically treated ejector nozzle was employed. A distinctive

operating feature is the independent control of core and fan (duct) stream temperature and exit velocity for

cycle matching with the following advantages:

• Meets thrust requirements over the entire flight spectrum

• Provides low-core exhaust velocity for noise benefits

• Minimizes cruise fuel consumption

At takeoff, the main burner is throttled to reduce jet noise. The duct burner is operated to provide an

inverted velocity profile. The thrust can be cut back for noise abatement after takeoff while still maintaining

the inverted velocity profile. During subsonic cruise, the VSCE operates as a moderate bypass turbofan engine

14
., i  ̧ ̧

with no duct augmentation. Variable geometry components are matched to reduce inlet spillage and bypass

losses. During supersonic cruise, bypass ratio and duct augmentation are reduced to maximize propulsive efficiency. Key parameters are:

• Overall pressure ratio

• Fan pressure ratio

• Design bypass ratio

• Maximum compressor discharge temperature

• Maximum turbine inlet temperature

A variable geometry bicone inlet with wing-leading

14.3

3.67

1.30

1,860°R

3,960°R

edge precompression was utilized based on the results

of Douglas supersonic transport studies conducted in the late 1970s. The nozzle concept is based on studies performed by Douglas and incorporates a combination of suppression techniques to meet FAR 36, Stage 3. Inlet

bleed air will be used for nozzle/engine cooling and then injected into the engine exhaust to aid in noise reduction. The engines are individually mounted in nacelles under the wing. Fuel will be supplied in shrouded lines

to the engine by routing it through the nacelle pylon. Pressure relief and drainage is provided. Maintenance

can be performed using ground access stands and conventional equipment.

The alternate Mach 3.2 engine is the GE variable cycle engine (VCE) using TSJE The VCE is twin spool,

with double bypass variable geometry to optimize fan/compressor/turbine match over the flight spectrum to

maximize subsonic and supersonic cruise performance. The engine is not augmented. The VCE incorporates

an axisymmetric nozzle with translating nozzle shroud and inner plug to vary the nozzle throat area. During

takeoff, the bypass flow is diverted through struts, forcing the flow along the inner plug to achieve an IVP for

jet noise reduction. Key parameters are:

• Overall pressure ratio 22

• Fan pressure ratio 4.8

• Bypass ratio 0.5

• Maximum rotor inlet temperature 4,000°F

The maximum temperatures the Mach 3.2 aircraft will experience range from 400 °F to 680°F (Figure 17).

The maximum temperature on the fuselage nose is 680°F; the leading edge temperatures are 625°F for the wing

and 530°F for the empennage. Fuselage lower surface temperatures range from 525°F to 590 °F, and the upper

fuselage surface temperatures range from 400°F to 460°E

Thermal management included both external and internal heat sources and dissipation requirements to

maintain energy balance. Fuel was used as the primary heat sink for absorbing energy loads. Except for the

engine inlet and nozzle, the aerodynamic heating of the aircraft was regulated by passive (i.e., insulation) thermal protection systems (TPS). Modularized Multilayer Insulation (MMLI), consisting of layered nickel reflector

foils was employed. The wing tank insulation limited the fuel tank temperature rise due to aerodynamic heating

to a maximum limit of 200°F for the TSJE The cabin air temperature was maintained at 70°E

Candidate structural materials consist of aluminum, titanium, polymetric composites (PC), aluminum

metal matrix composites (A1 MMC), and Ti metal matrix composites (Ti MMC). Up to 480°F, the A1 MMC

(SCS-8/RSR AI) has the best specific moduli value and the specific yield strength is relatively high. The Ti

MMCs are more suitable structural materials for higher speeds. The polymer composite material, Celion

6K/PMR-15, and RSR AI are possible candidates for the Mach 3.2 concept; however, they do not offer the

weight savings potential of AI MMC. Titanium materials having properties approximately 50 percent lower than

AI MMC would result in a less efficient and heavier structure. Buckling, crippling, stiffness, tension, and other

failure modes play important roles in the selection of materials for the primary structure of an aircraft with buckling and crippling accounting for approximately 75 percent of the primary structural weight. SCS-8/RSR AI is

the most efficient material considering buckling-crippling and thus was used for the structural material of the

Mach 3.2 concept.

Various flight conditions, maximum loads, maximum temperatures, and other pertinent conditions (flutter

and landing) were examined to verify structural integrity of the components. A finite element model (FEM)

of the Mach 3.2 concept was developed representing skins, spars, ribs, frames, and longerons. Fuel tank structure, landing-gear elements, nose/leading-edge structure, and LFC elements were included in the evaluation.

The FEM included the temperatures, materials, and loading conditions and was utilized to check stresses and

deflections to ensure structural integrity. Minimum gage materials were used as much as possible. The aircraft,

15
53O

400 *

460 *

\

540

m_ °F

680

59O

64O

53O

535

625

525

*UPPER BODY

FIGURE 17. MACH 3.2 MAXIMUM TEMPERATURES

under limit-load critical design case, was examined to ensure that there are no structural deformations or deflections that would inhibit movement of the control surfaces, such as the ailerons or flaps.

Various structural concepts for the major structural components were examined including conventional

skin-stringer design, super-plastically formed, diffusion bonded (SPF/DB) structure, and honeycomb construction. The optimization analysis included strength, stability, crippling, and thermal stresses. As a result,

SCS-8/RSR AI MMC honeycomb construction was selected for the load-carrying members. This resulted in a

structural weight reduction of 35 percent relative to mid-1970s technology studies of high-speed commercial

aircraft. The structures weights for the Mach 3.2 concept, including the wing, fuselage, and tail sections, were

derived from the structural optimization analysis based on commercial aircraft design criteria and standards.

Aircraft systems weights (excluding the TPS) were based on weight data from the Concorde and past

NASA/McDonnell Douglas high-speed commercial aircraft studies. System weights reflect assumed improvements due to lighter weight materials and innovative design. Advanced systems such as an all-electric vaporcycle cabin air conditioning and pressurization system, a fly-by-light flight guidance and control system, an allelectric secondary power system, and an all-electric anti-ice system are incorporated. The auxiliary power unit

system has been deleted with reliance on airport ground support utilities. The fuel system weight reflects integral fuel tanks. The LFC system weights include pumps, motors, ducting, controls, and the engine pneumatic

bleed system. In addition, an advanced technology glass cockpit (two-person crew) is assumed which includes

16
. 'c,, :J

flat panel displays and multiplexed digital data busses. The Mach 3.2 engine weight is based on data received

from P&W.

The operator items weight represents passenger and galley service to accommodate a three-class,

300-passenger cabin. The payload weight is derived as follows:

300 passengers at 165 pounds each

40 pounds of baggage per passenger

Mission payload weight

Cargo at 10 pounds per cubic foot

Maximum space-limited payload weight

= 49,500 lb

= 12.000 lb

= 61,500 lb

= 5.000 lb

= 66,500 lb

Mach 5.0. The Mach 5.0 concept, (Figure 18) features a highly swept delta planform with a buried internal

passenger cabin. The vehicle slenderness factor, Tau ( r = Vo1/Sprojected 3/2) is 0.069. The fully blended wingbody had a reference area of 17,000 square feet and an aspect ratio of 1.1. Inboard of the planform break, the

leading-edge sweep was 80 degrees and outboard the sweep was 60 degrees. Twin vertical tails are located outboard of the horizontal tip pitch/roll control surfaces. The single propulsion pod contains four GE variable cycle

hypersonic jet (VCHJ) engines. The propulsion system is highly integrated with the inlet and forebody and the

nozzle with the aft body. The landing gear has twin nose struts with dual wheels, and the main gear has two struts

with 12 wheels each. Liquid methane fuel tanks are arranged symmetrically around the center of gravity outboard of the pressurized cabin. The primary structure, cabin, and fuel tanks are structurally independent with

the pressure cabin suspended from longitudinal trusses.

t--- --- J_

, _.

1,'

LJ c: fl 11

E

I

-I

IOGW = 1,213(3OOPOUNDS

OEW = 473.805 POUNDS

S w : I 7.000 FI 2

I PAVtOA0

Fmsl crass = io PEnCENI = i Acnoss AI_21N m¢CH = _ sE_IS

Buslt_sS CrAss _ P_BCENI = 46 ACAOSS A! _ tN PllCH t0StAI$

COACHCtASS _Pt_CEN! = 4_*CROSSAI_2JN _ICN = ,BOStAIS

IOIAt 100PtAC_f_r = _O0S_AIS

FIGURE 18. MACH 5.0 CONFIGURATION

The fuselage was designed to accommodate a nominal seating arrangement of three classes: 10, 30, and

60 percent for first, business, and coach classes, respectively. The cabin cross section is a double-lobe type providing for twin aisles and eight-across coach seats, thereby minimizing frontal area and volume. Seat widths are

similar to MD-80/MD-11 size (i.e., 42-inch double-seats for coach class). The baseline interior arrangement

provides service for 300 passengers based on a maximum flight duration of 3 hours. Each class section has its

own lavatories, galley, coatroom, and cabin attendant stations. Three cabin doors per side lead to vertical entry/

exit chutes. No cabin windows are provided.

17 ORIGINAL PAGE IS

OF POOR QUALITY
The Mach 5.0 concept was based on previous NASA and industry studies (References 2 and 3) together

with detailed configuration integration through application of computational fluid dynamics (CFD) and

advanced graphics-based analysis. The blended-body, integrated engine/airframe concept was developed to

enhance both aerodynamic and propulsion performance. Laminar flow control was not included. The spatular

nose design, developed through CFD study, provided a blunted lift distribution which contributed to sonic boom

reduction (Reference 4).

The high-lift system consisted of plain trailing-edge flaps; there were no leading-edge devices. Longitudinal control and trim capabilities were achieved through elevons and movable tip controls. A nozzle flap was used

to enhance trim capability. Elevons and movable tip controls were used antisymmetrically to provide lateral control. Directional control was provided by rudders on the two vertical tails. A fuel management system was used

to aid longitudinal trimming.

The maximum trimmed lift-to-drag ratios for both overland flight profiles are shown in Figure 19. The

difference in the subsonic lift-to-drag ratios was due to flight at lower altitudes.

16

HA

E

(5

,<

rr

CI

b--

i1

J

12 _,,,,,__ SUPERSONIC FLIGHT PROFILE .OVER LAND

"_ SUBSONIC FLIGHT PROFILEOVER LAND

I I I I I

1 2 3 4 5 6

MACH NUMBER

FIGURE 19. MACH 5.0 LIFT/DRAG

The Mach 5.0 baseline engine is the GE variable cycle turbofan/ramjet engine, also referred to as a variable cycle hypersonic jet (VCHJ) using LNG fuel. The VCHJ (Figure 20) is a new engine concept defined by

GE in 1985 from NASA-sponsored supersonic transport studies during the late 1970s. The engine is an afterburning dual-rotor turbofan that combines double-bypass, variable-cycle engine features with high Mach flow

enhancement concepts. GE estimated a sideline noise of approximately 112 dB, or 10 dB in excess of the Stage 3

limit. Several engine operational modes are possible through the following features:

Low specific thrust at takeoff for noise reduction, with high-thrust capability through augmentation

Maximum climb and acceleration thrust during subsonic cruise and transonic operation

High-thrust windmilling (ramjet) operational modes plus efficient hypersonic cruise capabilities

Very good part-power subsonic cruise and loiter capabilities

The turbofan core engine of the VCHJ is based on the double-bypass, variable cycle engine (VCE) developed by GE during studies of supersonic propulsion technology. The VCE core employs a low-temperature

augmentor for supersonic acceleration. Noise constraints preclude the use of the augmentor at takeoff;

18
• "j

i:

ENGINE SHOWN WITH AXISYMMETRIC COANNULAR NOZZLE

FIGURE 20. GE MACH 5.0 VARIABLE CYCLE HYPERSONIC JET ENGINE

therefore, the fan has been sized for takeoff and subsonic cruise requirements without augmentation. The variable cycle features that give the VCE core improved flexibility over conventional mixed flow turbofans include:

• Split fan (outer) bypass duct with variable inlet guide vanes

• Fan variable area bypass injector

• Exhaust variable area bypass injector

• Variable area low-pressure turbine

• Core-driven rear fan block

• Variable area exhaust system with inve_ed velocity profile during takeoff

The turboramjet essentially phases out the turbomachinery during very high-speed operation. A ram air

bypass duct is located around the basic engine, and a special stream control valve similar to a GE patented variable area bypass injector (VABI) is employed to allow smooth transition from turbojet to ramjet mode above

Mach 3. Key parameters are:

• Overall pressure ratio

• Turbine rotor inlet temperature

• Bypass ratio

• Fan pressure ratio

25

4,000°F maximum

1.50

5.5

A variable geometry two-dimensional inlet was selected, both the inlet and nozzle are designed for

Mach 5.0 at an altitude of 83,000 feet. The four engines are mounted in a Quad Pod propulsion module on the

fuselage lower centerline. The engine installation is aerodynamically integrated with the fuselage. The Quad

Pod inlet includes movable ramps and a transition section with both bypass air and pressure surge dumps. A

single expansion ramp nozzle (SERN) has been incorporated allowing for a high degree of propulsion system

integration with contoured nozzle upper surfaces formed by the aircraft lower surface. The GE nozzle design

does not include an ejector due to the SERN, and hence it resembles a chute suppressor nozzle.

The engine compartment houses four separated engines and contains provisions for fuel lines, secondary

power-generating equipment, cooling air ducting, compartment ventilation, and fire protection. Engines will

be sealed to the inlet and nozzle ducts with flexible metal bellows.

The alternate Mach 5.0 engine is the Aerojet TechSystems two-spool air turboramjet (ATR) using LNG.

Aerojet initially proposed a conventional dual regenerator ATR where the incoming fuel is heated regeneratively in the combustor and nozzle walls to drive the turbine while providing cooling for these areas. The result

was an engine that yielded Mach 5.0 cruise performance equal to or better than that of the GE VCHJ. However,

the subsonic performance was significantly less than that of the VCHJ. Aerojet later proposed a two-spool ATR,

where a separate auxiliary airbreathing gas generator provides the motivating force for the turbine at flight

speeds below Mach 3.0. This results in significantly improved subsonic performance but at additional weight.

Although the latest Aerojet TechSystems data show potential for lower takeoff gross weights than for the VCHJ,

no comparative analyses were accomplished due to resource limitations.

The maximum temperatures the aircraft will experience range from 460°F to 1,580°F (Figure 21). They

are 1,580°F on the nose, 1,200°F on the leading edge of the wing, 1,425°F on the empennage, 920°F on the

upper fuselage, and 1,020°F on the lower fuselage.

19
1,000

1,010

1,010 *UPPER BODY

460'

1,040

1,425

960

FIGURE 21. MACH 5.0 MAXIMUM TEMPERATURES

Thermal management included both external and internal heat generation and dissipation requirements

to maintain energy balance. Fuel was used as the primary heat sink for absorbing energy loads. Except for the

engine inlet and nozzle, the aerodynamic heating of the aircraft was regulated by passive TPS employing MMLI.

The nonpressurized air gap external to the MMLI of the cabin and fuel tanks experiences low atmospheric pressure associated with flight at high altitudes. Consequently, the Mach 5.0 requires less insulation than the

Mach 3.2 concept. The cabin wall consists of a NOMEX honeycomb core with poly ether-ether ketone (PEEK)

face sheets as the pressure shell. The cabin insulation criteria included a ?0°F cabin air temperature, an 80°F

cabin liner temperature. The MMLI was optimized for the minimum total weight considering LNG boiloff, insulation, and tank structure. The throat regions of the Mach 5.0 nozzle and inlet require active cooling. Based

on McDonnell Douglas studies, a wall structure is recommended comprised of honeycomb and a skin heat

exchanger which serves to transfer heat from inlet air and engine exhaust to the fuel.

Candidate materials consist of titanium (including MMC), Ren6 41, and TD Ni Cr. SCS-6/RSR-Ti, being

the most efficient considering specific moduli and specific compressive yield strength, was used for the structural

material of the Mach 5.0 concept. Ren6 41 has the highest temperature capability and was used for the wing

leading edge. The TD Ni Cr was used for the leading edge of the empennage because of its greater temperature

overshoot capability. Considerable work is being done in industry and government, especially NAS,_s Langley

Research Center and Lewis Research Center, to develop these and other advanced materials. Where TD Ni Cr

is no longer offered commercially it would only be considered if the R&T program was not successful. McDonnell Douglas has performed high-temperature tests on titanium MMC with success in the Mach 5.0 temperature

range.

In addition to the materials for the major components, materials investigations were performed for the

major structural elements of the propulsion system, including the inlet ramps, nacelle, exhaust, and exhaust

ramps. The Lewis Research Center has developed a ceramic-ceramic material far superior to others. Consequently, it was chosen for the internal structural applications. The external shell of the inlet and nacelles is

SCS-6/RST Ti, and Graphite/HFC,TZM was used for the engine/nozzle transition structure.

The same type of loading conditions were examined for the Mach 5.0 concept as for the Mach 3.2 concept.

In this case, however, the cruise condition was the critical design case. An FEM was constructed for the Mach 5.0

concept, representing all the structural elements of the aircraft: the spars, ribs, frames, longerons, and shear

20
websforthemajorcomponents. Thepassenger cabinandpropulsionsystem,includinginlet,nacelle,and

exhaustalsoaremodeled.TheFEMincludedtemperatures, materials,andloadingconditions.Bothdeflection

andstresswerecheckedtoensurestructuralintegrity.Stresslevels,aswellasdeflections, werewithinstructural

allowables.

As in the case of the Mach 3.2 evaluation, several structural concepts were examined resulting in the three

best choices; (1) skin-stringer, (2) superplastic formed/diffuser bonded, and (3) honeycomb construction. The

selection and optimization of the SCS-6/RSR Ti honeycomb construction resulted in a structural weight reduction of 15 percent relative to the mid-1970s technology studies of high-speed commercial aircraft. The structures

weights for the Mach 5.0 concept, including the wing and movable tip, fuselage, and tail sections, were derived

from optimization analysis-based commercial aircraft design criteria and standards.

Aircraft systems weights (excluding the TPS and the liquid methane fuel tank) were based on weight data

from the Concorde and past NASA/McDonnell Douglas high-speed commercial aircraft studies.

The Mach 5.0 concept system weights reflects assumed improvements due to lighter weight materials and

innovative design. The weight of the LNG fuel tank structure is based on optimized structural analysis.

Advanced systems such as an all-electric vapor cycle cabin air conditioning and pressurization system, a fly-bylight flight guidance and control system, an all-electric secondary power system, and an all-electric anti-ice system are incorporated. The auxiliary power unit system has been deleted with reliance on airport ground support

utilities. In addition, an advanced technology glass cockpit (two-person crew) is assumed, which includes flat

panel displays and multiplexed digital data busses. The Mach 5.0 engine weight is based on data received from

GE.

The operator items weight represents passenger and galley service to accommodate a three-class,

300-passenger cabin. The payload weight is as follows:

• 300 passengers at 165 pounds each = 49,500 lb

• 40 pounds of baggage per passenger = 12,000 lb

• Mission payload weight = 61,500 lb

• Cargo at 10 pounds per cubic foot = 5,000 lb

° Maximum space-limited payload weight = 66,500 lb

i •

21
22
....i_i

PERFORMANCE AND ECONOMIC S

Performance

The HSCT performance was analyzed according to commercial domestic and international aviation rules

and practices. The mission profile is depicted in Figure 22 and begins with conventional takeoff and climb out

to 10,000-feet altitude. This is followed by an accelerating climb to the cruise Mach number. The climb is continued at cruise Mach number until optimum cruise altitude is reached. The main limitations in climb are the cabin

rate of climb and the aircraft excess thrust over drag at the top of climb. Conventional cabin pressure altitude

at cruise is 8,000 feet, and the limiting rate of pressurization change is equivalent to 300 feet per minute at sea

level. This requires that the climb takes at least 23.5 minutes. A requirement of 4,000-feet-per-minute potential

rate of climb was assumed to ensure sufficient acceleration and rate of climb to reach cruise altitude.

SUPERSONIC

CRUISE CLIMB

1.7 MIN ]

TAKEOFF / CLIMB/ACCEL TO:

I ! M 1.0 AT 30,000 Fq" /

RANGE

SUBSONIC

LONG-RANGE CRUISE

4 MIN AT 30,000 FT"

APPROACH

6MIN / / 30MINAT

I I 5"/,, /

II Ii BLOCK ]

[ [ t FUEL /

i/

200 N MI

1 I1"_ TO ALTERNATE

AIRPORT

FIGURE 22. MISSION PROFILE

Cruise is flown at constant Mach number and optimum altitude to maximize range. During fuel burn-off

the aircraft is allowed to cruise-climb to remain at optimum conditions. The descent is at idle power and constant

airspeed, as is the current convention. The cabin rate of descent is limited to 300 feet per minute with idle power.

Below 10,000-feet altitude, regulation-specified speeds of 250 knots are maintained until landing approach

at conventional speeds of 140 knots or less. Fuel reserves based on international rules are maintained: 5 percent

of block fuel, fuel to fly to an alternate destination of 200 nautical miles, and fuel for a half-hour hold at the

2,000-foot altitude. Sufficient taxi, takeoff, and landing allowances for time and fuel determinations are

included.

Winds aloft at the 60,000- to 90,000-foot cruise altitudes of the Mach 3.2 and Mach 5.0 concepts are considerably less than those at the cruise altitudes of subsonic aircraft (35,000 to 43,000 feet). For supersonic vehicles,

wind influence on still-air range is less than a tenth of that of subsonic airplanes. Therefore, wind effects are

only used for the subsonic reference airplane. The effects of earth's rotation are incorporated in the mission

analysis.

Concept sizing was accomplished by varying wing reference area and engine size, and taking into account

constraints and margins as appropriate. These constraints are: (1) takeoff field length of 11,000 feet or less,

(2) landing approach speed of 140 knots or less, and (3) cruise at optimum altitude or at the operationally determined ceiling. The minimum value of design takeoff gross weight satisfying these constraints determined the

wing/engine size combinations.

The economic analysis included an advanced subsonic transport with year 2000 technology. The airplane

was sized for 300 passengers and 7,400-nautical-mile still-air range. (The design range of 6,500-nautical-mile

geometric distance was corrected for prevailing westerly headwinds across the northern Pacific.) The wing and

PRECEDING PAGE BLANK NOT FILMED

23
engines were sized by an initial cruise altitude for maximum range and cruise-power ceiling altitude. The maximum takeoff gross weight is 397,000 pounds, with a takeoff field length of approximately 7,000 feet and landing

approach speed of 130 knots.

Due to operational uncertainties regarding overland operations, two configurations were developed for

Mach 3.2. The D3.2-3A configuration was optimized for supersonic flight with minor compromises for lowspeed flight (takeoff and landing). The D3.2-4B configuration was compromised for subsonic flight over land.

For the 6,500-nautical-mile design mission, the -3A was sized at a maximum takeoff gross weight of 769,000

pounds, and the -4B was sized at an MTOGW of 880,000 pounds. A representative city-pair was selected to illustrate the impact of disallowing supersonic flight over land. The New York-Tokyo market is second largest in

terms of international revenue passenger miles (rpm) and thus presents an aircraft sizing focal point. If supersonic flight over land could be achieved, the HSCT great circle distance would be 5,846 nautical miles. If supersonic flight over land is disallowed, a different routing would be expected to minimize flight time by reducing

the distance over land from 4,526 nautical miles to 2,518 nautical miles for the subsonic operation. This route

has an overall distance of 6,236 nautical miles; the subsonic flight segment represents 40 percent of the total

flight if the HSCT is diverted from the great circle route to achieve minimum flight over land (Figure 23).

K

GREAT 600 720 4,526

A I

CIRCLE -

WATER

[ TOK,'OI

3,718

REROUTE %_/ / _///i /////////////

WATER

2,518

SUBSONIC

[ NEW YORK]

5,846 N M I

FIGURE 23. KEY OVERLAND MARKET NEW YORK-TOKYO

- 6,236 N MI

To illustrate the effect of subsonic flight segments, the -3A and the -4B concepts were sized to

6,500-nautical-mile range with up to 40-percent subsonic legs. Figure 24 shows the increased maximum takeoff

gross weight for the -4B concept for the all-supersonic design mission. This is a result of greater structural weight

due to the larger aspect ratio and the lower aerodynamic efficiency at supersonic cruise. However, with increasing subsonic flight distances, the supersonic design is penalized with increased maximum takeoff gross weight

as contrasted to the improved subsonic configuration. Consequently, for subsonic legs greater than 27 percent

of the total range (approximately 1,750 nautical miles), the -4B concept is lighter than the -3A configuration.

The best range of the Mach 5.0 concept was 3,900 nautical miles with a maximum takeoff gross weight of

1,213,000 pounds (meeting the takeoff field length requirement of 11,000 feet). With this limited range capability, an economic comparison with other vehicles is not possible; therefore, technology improvements are

assumed to produce a vehicle capable of a 6,500-nautical-mile range. Early Mach 5.0 configuration studies

showed a promising lift-to-drag ratio of 6.25 compared to 5.4. Further configuration development is expected

to lead to lift-to-drag levels of 6.75. This corresponds to a 20-percent decrease in drag. Further, considering that

24
.x

MTOGW

(1,000 LB)

950

900

85O

8OO

750

-- -4B _/

780 1,750 2,200

II I i I 1

0 10 20 30

SUBSONIC RANGE

2,600 _ NAUTICAL MILES

I

40 _ % OF 6,500 N MI

FIGURE 24. MACH 3.2 CONCEPT SIZINGS FOR PARTIALLY SUBSONIC RANGE

the Mach 5.0 engine efficiency is 13 percent below the potential level, a 10-percent decrease in fuel flow was

applied. These technology advancements would have later technology readiness dates. With these improvements in drag and fuel efficiency, the Mach 5.0 concept with a takeoff gross weight of 1,213,000 pounds achieves

the 6,500-nautical-mile design mission with a takeoff field length of 8,600 feet. In retrospect, Mach 5.0 can be

achieved with LNG fuel (heat sink potential for the greater aerodynamic heating compared to Mach 3.2); however, the energy level of LNG is not adequate to meet performance requirements.

Weight comparison of the Mach 3.2 and 5.0 baseline concepts is shown in Figure 25. The dominating

weight fraction is the fuel. To reduce the takeoff gross weight and make the HSCT more viable and competitive,

the fuel fraction must be reduced considerably. Future HSCT studies should address innovations in drag reduction, propulsion system inlet and nozzle design, engine cycle, and structural concepts providing safe, maintainable, and reliable components that produce reduced fuel fractions.

Figure 26 depicts the overall fuel efficiency in terms of pounds of fuel per available seat-nautical mile to

enable comparison with other airplanes, such as the 108-seat, Mach 2.0 Concorde. The Mach 3.2 concept,

through year 2000 technology, would be expected to provide a 50-percent reduction in fuel requirements per

seat compared to Concorde. The values given are 3,500 nautical miles as a representative system average length

of all-supersonic range and of the split half-subsonic, half-supersonic flight. The Concorde data are for the

3,000-nautical-mile range. Values for the DC-10 (277 seats) and 747 (365 seats) are shown as 1970's subsonic

data.

Economics

Figure 27 details the commercial value assessment procedure developed during this study. Interactions

between HSCT characteristics, market parameters, economics, and competitive factors are noted in simplified

format. Basic assumptions include: (1) supersonic flight over land is disallowed, (2) passenger market is segmented into four fare classes, (3) seating is based on the onboard passenger mix, and (4) fuel price is station-specific. All costs are expressed in 1987 dollars.

The traffic model under the speed over land restriction consists of eight regions -- a subset of the unrestricted model. Europe-to-Africa and Europe-to-the-Far-East were eliminated because all routes involved long

subsonic distances. Analysis of a Tokyo-based airline indicates a 10- to 15-percent productivity loss due to the

overland restriction. Similar analysis of a Paris-based system showed 7- to 28-percent losses. These results are

25
PAYLOAD 8.9%

SYSTEMS 11.8%

POWER PLANT 4.6%

I I I,dl_====_

STRUCTURE

12.8%

MACH 3.2 (-3A)

TOGW = 769,000 LB

POWER PLANT 16.5%

SYSTEMS 10.2% PAYLOAD 5.6%

:'" STRUCTURE

11.8%

 iiiiiiiiiiiiiiii!' iiiiii iiiiiiii ; .:

MACH 5.0

TOGW= 1,213,000 LB

0.4

0.3 u

LB FUEL

0.2 --

ASNM

01 --

0

CONFIGURATION

SEAT

FIGURE 25. WEIGHT BREAKDOWN

Fro.

I

I

I

I

I

1

1970S CONCORDE SUBS REF

S UBS 108 300

STAGE LENGTH 3,500 N MI

mm_ mi_

I i

MACH 3.2(-3A) MACH 3.2(-4B) MACH 5.0

300 300 300

/I I. ALL /4

/ / SUPERSONIC

//

//

I/

I/

II

FIGURE 26. FUEL EFFICIENCY

corroborated by MIT, indicating an 8- to 20-percent loss in productivity. Productivities from the Douglas computer model show a 10- to 15-percent loss compared to the potential supersonic flight over land. The speed

restriction reduces productivity, which reflects two significant factors. First, city-pairs with long distances over

land may not be flown in the restrict.ed case because they exceed the range capability (reduced because of the

subsonic cruise requirement). These city-pairs are precisely the ones for which the potential penalty is greatest.

Secondly, the restriction causes rerouting to avoid land masses. Although the diverted flight path is longer, it

26
i,

I SUBSONIC '

BLOCK TIMES

I BLOCKHSCTTIMES

FARE PREMIUML_

o,.1,.2,.3.... I'1

FARE DATA

ADVANCED

SUBSONIC

ECONOMICS

TIME SAVINGS

HSCT FARES

BY FARE CLASS

H SUBSONIC BY FARE CLASS FARES

T

ISCHEDULING

EXERCISE H

I

FLEET SIZE

l FREQUENCIES ANNUAL

_ [ AIRCRAFT PRICE H ECONOMIC BENEFITS I

}__ CASH

OPERATING

COSTS

ASM DEMAND 1

I FACTORS LOAD

HSCT

CAPTURE PASSENGER

CAPTURE

HA,RO.AFTI iP"OGRA

wo.THI "[V'AB'"TY I

, 11

I I CONFIGURATION SELECT i I REVENUE ANNUAL I

H PASSENGER M,x , VE'TN

FIGURE 27. ECONOMIC AND MARKET STUDY MODEL

is flown at high speed and the flight time is favorable compared to that of the great circle, subsonic flight paths.

Finally, whether aircraft cruise over land at subsonic speed or fly a longer path to avoid land or a combination

of the two, blocktime is certain to increase. Each city-pair (no flights under 2,000 statute miles) in the eightregion model was studied to estimate the length of subsonic and supersonic cruise legs and to devise a better

route that minimizes overland travel. City-pairs under 2,000 statute miles tend not to be very good HSCT candidates, because block-time savings over subsonic aircraft are not significant at short ranges.

Forecast scenario 2 traffic levels were assumed. The North Atlantic and Pacific routes are the most important with the Pacific becoming the greater after the year 2000. During the period 2000 and 2025, the projected

growth in the Pacific will be more than twice that of the Atlantic. Thus, the design range should be established

so that the HSCT should be able to perform most of the transpacific segments. A range of 6,500 nautical miles

will permit flying from Sydney to the west coast of North America. Ten-percent traffic stimulation is assumed,

(i.e., 10percent of the dollar value of the passenger's time savings, excluding fare premium, is used to purchase

additional trips). The advanced subsonic airplane is in competition with the HSCT in terms of market capture.

The operating cost components follow CAB Form 41 format for direct and indirect cash costs. These are:

flying operations, maintenance, passenger service, aircraft and traffic servicing, promotion and sales, and general and administrative. Cost estimates are computed by Douglas operating cost formulas. Information generated during the study includes fuel costs and fuel infrastructure costs, and maintenance cost. Fuel costs

(i.e., feed stock prices) were determined through industry-wide workshops. Fuel handling costs, including

special facilities where required, resulted from fuel facility sizing and costing studies performed by engineering

companies experienced in such projects. The results are shown in Figure 28. Airframe maintenance cost

predictions utilized labor-hour and material cost data for eight large, commercial transport aircraft and the estimated effect of future high-speed civil transport technologies. Airframe maintenance labor-hour and material

cost data were analyzed to determine the most maintenance-intensive operating systems by ATA chapters. Concorde maintenance information was used subjectively due to its unique operating environment; small fleet size,

limited support, high occurrence of unique operations (charters) , spare parts available, and limited technical

27
°

30

FUEL

COST

(CENTS/

PER

POUND)

20

10

FEED STOCK PRICE

FUEL FACILITY COST

HIGH (.... ) TO LOW

VOLUME FACILITY

NOTED

9.1

p41111,,' II i,III iiIII ijlII illII illII ip _ i,IIIi,, _

JET A

13.2

21.1

5.7

' ,,_111 _i1_1_ _11_11_1__ _11_' ,_ III iiii_1ip__,_'_ iii

LNG

FIGURE 28. FUEL COST

support. The X-15 and SR-71 were analyzed to determine peculiar support requirements. Engine maintenance

costs were derived by Douglas from data provided by P&W and GE.

Figure 29 shows the cash operating costs for the Mach 3.2 and Mach 5.0 aircraft and for the advanced subsonic aircraft of the same size. Ownership-related expenses are not included because cash flow over the life of

the HSCT is used to compute its value as an investment. The Mach 3.2 aircraft cash operating costs on a per

seat-mile basis are estimated to be approximately 30 percent higher than for the subsonic transport. The

Mach 5.0 aircraft is estimated to have a cash operating cost approximately 140 percent higher than the subsonic

transport. Comparison of the DOC per seat-mile to the hourly costs highlights the value of speed and underscored the need to achieve supersonic cruise overland operations.

6o

50

4o

DOLLARS

PER BLOCK

HOUR 30

(THOUSAND)

20

10

NO OWNERSHIP-RELATED COSTS

FLYING OPERATIONS

(FUEL) 5O,200

22,200

MACH 3.2 MACH 5.0

(-3A).

7,400

ADVANCED

SUBSONIC

12

10

CENTS PER 8

SEATSTATUTE 6

MILE

4

2

0

4.5

ADVANCED

SUBSONIC

10.7

ii!!!i!iiiiiiiiiiiiiiiiiiiii

....................

...................

...................

..................

...................

:::::::::::::::::::::::::::::

5.8 i_i!i!i!i?_i_!_iii_ii_iiiil

:::::::::::::::::::::::::::

_iiiiiiiiiiiiiiiiiiiii!i!!!!i

iiiiiiiii_!_!_!_!!!!!!!!!!_i

_iiiiiiiiiiiiiiiiiiii!ii_i

:.:-:-:.:.:.:.:.:.:.:.:.:.:!

iiiiiii!i!:!:!:!:!:?:i:!:i:!

MACH 3.2 MACH 5.0

(-3A)

FIGURE 29. AIRCRAFT OPERATING COSTS

28
• i_ ,

V_

Passenger revenue is based on published International Civil Aviation Organization (ICAO) fare data, fare

premium assumptions, and corresponding HSCT market share statistics. The passenger market is assumed to

consist of equal numbers of business and personal travelers. Both the business and personal components are

segmented by fare classes. Each fare class has its own average value of time and fare level (related to ICAO

full fares). As fare premium increases, the HSCT market share is reduced. Almost all of the fare-sensitive traffic

will be lost to subsonic aircraft at 10- to 20-percent fare premium while 50 percent of the first-class and businessclass passengers would remain with the HSCT at 30-percent fare premium. Yield (cents per passenger-mile)

is improved because fares increase and the onboard passenger mix changes to favor the higher yield business

and first-class passengers.

Effective yields are significantly different from what might be predicted from load factors and published

fare. Frequent flyers, upgrades to first class, nonrevenue passengers, and special low fares from interline agreements cause the effective yield to be diluted and significantly lower than that predicted from load factors and

ICAO fares. The revenue based on ICAO fares is adjusted such that an advanced MD-11 would have an investment value (aircraft worth) equal to its selling price. This ensures that the HSCT revenues are based on realworld yields and that HSCT aircraft worth estimates are consistent with the market price of a known airplane.

Figure 30 illustrates the annual revenue-generating capabilities of the Mach 3.2 and Mach 5.0 concepts with

the Mach 5.0 concept generating approximately 10 percent greater revenue.

DOLLARS

(MILLION)

25O

2OO

150

100

50

m

MACH 3.2 (-3A)

o I I I I

0 1 0 20 30 40 50

FAREPREMIUM (PERCENT)

FIGURE 30. ANNUAL REVENUE

Aircraft worth is the investment value of an airplane to the airline operator. The worth of an HSCT is

estimated by a process that determines the price to the operator (i.e., airline) so that a target rate of return on

investment is achieved by the operator. This process includes 1987 tax law and depreciation schedules, life of

the asset, and, most importantly, the annual operating cash flow. All of the airplane characteristi c such as size,

weight, speed, lift-to-drag ratio, propulsion efficiency, and other parameters are embodied in the cash-flow estimates. Also involved in the cash flow (and hence, aircraft worth) are operational parameters such as utilization,

turnaround time, passenger mix, load factor, fare differences in various regions of the world, and fare premium.

29
ResultsareshowninFigure31highlightingthesignificantadvantageestimatedfortheMach3.2aircraftcomparedtotheMach5.0aircraft-- $200 million turnaround.

Aircraft prices were developed using the total McDonnell Douglas base of experience and knowledge in

the field of high-speed technology and support efforts (e.g., materials, processes, and the like) to the maximum

extent. Labor and material resources were estimated on a discrete evaluation basis. Resources were estimated

by major aircraft system/component and by functional category. Figure 32 illustrates the Mach 3.2 concept production breakdown. Development costs included all of the necessary resources and tasks required to design,

develop, produce, and demonstrate an aircraft that can be FAA certified. Labor hours were translated into constant 1987 dollars using the aerospace fully-burdened labor rates for the different categories of labor (e.g., Engineering, Tooling, Quality Assurance, and so on). Material and equipment were estimated separately with propulsion system costs furnished by GE and P&W. The end product of the estimating process is a flyaway cost

(price) in which the development cost is amortized over the production program including the manufacturer's

targeted rate of return.

20

10

DOLLARS

(MILLIONS)

0

-10

ANNUAL CASH FLOW

.i!!iiiiiiiiiiiiiiiili!iiiilililililililiiiiiiiiiiiiiiiii

//

-14

150

100

DOLLARS 50

(MILLIONS)

-50

-20 -100

AIRCRAFT WORTH

114

....................,........,................................

:i:_:_:_:_:_:_:_:_:_:_:_:_:_:_:i:i:i:i:i:i:i:i:i:i:i:i:i:i:!:i:

.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:

:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.

:::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::

............................................

===============================================================

+:+:+:+:+:+:+:+:+:+:+:+:+:+:+:+

MACH 3.2 (-3A)

MACH 5.0

FIGURE 31. HSCT ECONOMICS AT ZERO FARE PREMIUM

Necessary conditions for economic viability include: (1) airplane revenues covering operating costs plus

an attractive rate of return to the operator, (2) fares low enough to provide competitive HSCT service, and (3) a

market large enough to permit a selling price lower than the investment value of the airplane. The evaluation

procedure places the HSCT in competition against the advanced subsonic airplane on a city-pair basis. This

ensures that the HSCT is applied to those markets in which it performs best. Repetition of this procedure for

various fare levels will determine whether the foregoing conditions can be simultaneously satisfied. Production

to satisfy the market for the period 2000 to 2025 ranges from over 1,500 (zero fare premium) to 250 units (at

approximately 40-percent fare premium). The Mach 3.2 aircraft becomes economically viable at 25- to 35-percent fare premium. Results are shown in Figure 33.

The Mach 3.2 (-3A) concept is potentially viable with a fare premium of 30 percent (500 units). These data

show that the Mach 5.0 concept is not economically viable at any fare premium. Figure 34 shows potential

increases in aircraft worth based on technology improvements to both airframe and engine as well as operational

improvements through supersonic overland flight and reduced turnaround time. This, together with a modest

fare premium, increases aircraft worth to approximately match the estimated Mach 3.2 concept flyaway price.

The HSCT promises economic benefits including higher levels of U.S. gross national product (GNP), a

better balance of trade, and increased employment. The size of these benefits depends upon the size and timing

of the effort from research through production and delivery. They also depend upon the ability of the economy

to absorb the increased spending. Generation of great magnitudes of spending over too short a period of time

will lead to displacement of other business investment (crowding out), higher interest rates, and increased inflation. A moderately sized program, however, will generate favorable changes in the economy. Historically, for

30
FIGURE 32. MLCH 3.2 CONCEPT PRODUCTION BREAKDOWN

8OO

600

O

400

co

n-

_J

0 200

a

MACH3.2 MACH5.0

FLY AWAY PRICE

__ 1500 1

@ -- _ PRODUCTION

QUANITY

-- / I ECONOMIq

I VIABILITY I

"'--- AIRCRAFT WORTH _

I I I I

0 10 20 30 40 50

FARE PREMIUM (PERCENTAGE)

800

0

•_ 1500

,,_t

4oo _

Or)

n20o

0

0

PRODUCTION QUA_

"-- FLY AWAY PRICE

AIRCRA_

I ,,-'f I I I

10 20 3O 40 50

FARE PREMIUM (PERCENTAGE)

FIGURE 33. HSCT WORTH/PRICE COMPARISON

every dollar of direct expenditures in the U.S. aerospace industry, approximately $1.20 of indirect gross national

production occurs. This results in a total effect of $2.20 on GNP for every $1.00 spent on aerospace. The multiplier of 2.2 incorporates a normal amount of crowding out. For every new aerospace job created by the HSCT,

there will be other indirect jobs created. It is estimated that each direct aerospace job will lead to about 1.5

indirect jobs resulting in an employment multiplier of 2.5. There can be a crowding out effect in employment;

if the unemployment rate becomes too low, then wages will be bid up and inflation and unfilled jobs may result.

None of the scenarios considered here, however, seem likely to present significant dangers of employment

crowding out, given a baseline prediction of 5-percent unemployment in 2013.

Several of the key parameters of the economic benefits analysis have been identified. In order to bound

the problem numerically, a range of values is assumed for some of the key parameters: (1) total production is

31
POTENTIAL IMPROVEMENTS

• GROSS WEIGHT REDUCTION: $52M

• IMPROVED ENGINE EFFICIENCY: $28M

• JET "A" FUEL SPECIFICATION: $35M

• SUPERSONIC CRUISE OVERLAND: $50M

• 1-HOUR TURNAROUND: $35M

• HIGHER COACH/DISCOUNT COACH RATIO: $45M

DOLLARS

(MILLIONS)

800

600

400

200

._FLY-AWAY PRICE 500

:$so_\\ ....... /250

PRODUCTION

OUANT,TY

_ AIRCRAFT WORTH

I I I I

10 20 30 40 50

FARE PREMIUM -- PERCENTAGE

FIGURE 34. MACH 3.2 ECONOMIC VIABILITY

500 and 1,000 aircraft; (2) unit production cost is $322 million and $301 million; (3) development cost is $14.4

billion; (4) domestic content is 50 percent, 70 percent, and 85 percent of production cost (Figure 35). Also, it

is assumed based on past U.S. commercial aircraft programs that 73 percent of the aircraft sold are exported.

Assuming 70-percent U.S. production, a 1,000-unit program would create more than 400,000 jobs at peak employment and produce a gain in GNP of nearly$500 billion, resulting in more than a $100 billion improvement

in the balance of trade. The peak effect on GNP occurs in 2013 when production levels first reach their highest

level. The sum of the improvement in the U.S. balance of trade increases very sharply with the percent of U.S.

production.

Very large programs with effects in excess of 0.5 percent of GNP may dampen the growth through crowding

out and result in multipliers considerably less than 2.2. Scenarios 5 and 6 fall into this category. All of the scenarios show an improvement in balance of trade. In perspective, so long as 30 percent or more of the production

of the HSCT is performed in the U.S., the project will improve the balance of trade. The analysis of the balance

of trade effects treats only direct effects of importing and exporting.

Technology Availability Dates

As stated previously, this study assumes an HSCT certification date of 2000/2010 with a TAD of 1995/2000.

An overall perspective of technology effects on HSCT gross weight relative to 1990 (current) technology is presented in Figure 36. Ten-year technology advancements (year 2000 versus 1990) are projected to result in a

27-percent TOGW reduction. Of this, 13 percent is attributed to the utilization of aluminum metal matrix composite structure (ALMMC), a 5-percent improvement from the application of a partial wing LFC system for

drag reduction improvement, a 6-percent improvement reflecting improvements in engine thrust-to-weight ratio and improved SFC, and a 2-percent improvemen_ for aircraft systems technology weight improvement. Figure 36 shows a further takeoff gross Weight reduction of 20 percent based on year 2010 technology; a full chord

LFC system accounts for an additional 10-percent weight improvement, 6 percent would be provided by innovative improvement in engine T/W and SFC, and approximately 4-percent weight reduction is achieved by innovation in aircraft systems and structural design.

Identification of key propulsion-related technologies that affect the HSCT through the potential for

achieving HSCT economic viability and environmental acceptability was a key element of this study. All three

engine companies (GE, P&W, and Aerojet TechSystems) provided inputs and projections for future improvements. For the baseline, an FAA certification date of 2005/2010 was assumed, with projections beyond that to

an FAA certification date of 2015/2020. Technology areas that could result in reduced TOGW include improved

cycle performance, both at subsonic and supersonic cruise, and reduced engine weight (increased thrust/weight).

The two items are directly related to economic viability.

32
L

.i

/,

SCENARIO

1

I

R&T EXPENOITURES ($ BILLION)

UNITS PRODUCED

PROOUCTION COST PER AIRPLANE

($ MILLION)

PERCENT U.S. PRODUCTION

PEAK EMPLOYMENT GAIN

PEAK DECREASE IN UNEMPLOYMENT

RATE (FROM BASE RATE OF

5 PERCENT)

TOTAL GNP GAINOVER PROGRAM

LIFE ($ BILLION)

PEAK GNP GAIN -- YEAR 2013

($ BILLION)

PERCENT INCREASE IN GNP

-- YEAR 2013

TOTAL IMPROVEMENT IN BALANCE

OF TRAOE OVER PROGRAM LIFE

($ BILLION)

50

167,000

0.11

$2O9

$14

0.17

$37

SCENARIO

2

$14.4

5O0

$322

70

234,000

0.16

$280

SCENARIO

3

85

284,000

0.19

$333

SCENARIO

4

5O

312;500

0.21

$363

$20

SCENARIO

5

$14.4

- 1,000

-- $301

7O

437.500

0.29

$495

SCENARIO

6

85

531,000

0.35

$595

0.23

$69

$24

0.28

$92

$26

0.31

$69

FIGURE 35. MACH 3.2 HSCT BENEFITS TO U.S. ECONOMY

$37

0.44

$129

$45

0.53

$175

P&W conducted a subsonic performance improvement study for the VSCE in terms of the major cycle

parameters -- fan pressure ratio, bypass ratio, overall pressure ratio, and combustor exit temperature. Study

results showed that increasing overall pressure ratio offered the best potential for improvement in subsonic SFC

but the improvement was only 1.4 percent. GE performed a similar study for its Mach 3.2 VCE, examining the

effects of control schedules optimized for Mach 0.95 cruise. This indicated a potential reduction in SFC from

0.5 to 2.0 percent. It is concluded that other approaches must be taken to reduce aircraft subsonic fuel

consumption.

Both engine companies were requested to provide projections of SFC improvements assuming better

technology. These improvements would be achieved through component efficiency increases, reductions in

cooling air requirements, improvements in burner efficiency, and reductions in nozzle unburned air. Increased

material temperature capability to reduced losses and better heat transfer formed the bases for these projections

to increase engine thrust/weight, thus reducing TOGW. Projected SFC reductions (5 percent subsonically,

3 percent at Mach 3.2) would result in an approximate 6-percent TOGW reduction for the all-supersonic cruise

mission and ll-percent TOGW reduction for the mission with 2,000-nautical-mile subsonic cruise segment over

land.

P&W has established a goal of a 10- to 15-percent improvement in engine T/W based on three major

improvement areas:

• Materials technology • Turbine technology

• Cooling system technology

An additional 10-percent improvement would be achieved if the design jet velocity could be increased to 2,800

feet per second and still satisfy FAR Part 36, Stage 3 noise requirements with no additional nozzle weight

penalty. Combining these two projections and including an allowance for additional noise-suppression devices,

33
i_I_I:_

z

100

90

8O

PERCENT

TOGW,

FROM

1990

TECHNOLOGY

70

60

5O

PROPULSION

PROPULSION

MATERIALS

AERODYNAMICS

SYSTEMS

MACH 3.2 HSCT

TITANIUM

ALMMC

2000

TECHNOLOGY

(ALMMC)

2010

TECHNOLOGY

FIGURE 36. PROJECTION OF POTENTIAL WEIGHT REDUCTION

an overall increased thrust/weight of 20 to 25 percent appears reasonable. The corresponding TOGW reductions

would be:

• 9 to 11 percent for the all-supersonic cruise mission

• 14 to 17 percent for the mission with 2,000-nautical-mile subsonic cruise segment overland

Although engine weight is only approximately 10 percent of TOGW, there is the potential for TOGW reductions

of 10 percent or more from increasing engine thrust/weight. When combined with the SFC reduction, the potential TOGW reductions would be:

• 15 to 17 percent for the all-supersonic cruise mission

• 25 to 28 percent for the case with 2,000-nautical-mile subsonic cruise segment overland

GE and Aerojet TechSystems have addressed the technology needs and projections for their respective

Mach 5.0 engines.

GE showed an SFC improvement of 0.5 percent for its variable cycle turbofan ramjet engine through control schedule optimization. GE provided projections of SFC improvements (5 percent subsonically, 2 percent

at Mach 5.0) assuming better technology. These would be achieved through component efficiency increases,

reductions in cooling air requirements, improvements in burner efficiency, and reductions in nozzle unburned

air. Increased material temperature capability, reduced losses, and better heat transfer formed the bases for

these projections. Projections of 20- to 25-percent improvement in Mach 5.0 engine thrust/weight appear reasonable through higher stage loadings and advanced materials. When combined, these data indicate potential

TOGW reductions of:

16 to 18 percent for the all-supersonic cruise mission

33 to 36 percent for the case with 2,000-nautical-mile subsonic cruise segment overland

34
AIRPORTS AND THE ENVIRONMENT

Airports

Most, if not all of the world's international airports that are likely candidates for future HSCT service

already experience DC-10 and 747 aircraft with larger versions forthcoming. Few new airports are expected in

the future for several reasons. First, the very high cost; the new Osaka Airport is reported to cost in excess of

$5 billion and the new Denver Airport at more than $2.5 billion. Second, the lack of public acceptance of new

airports causes significant delays and has forced development in remote locations, less convenient for the

traveler. And third, the flatter terrain is already developed leaving topographic areas inappropriate for airport

development. For these reasons, new airports are not good solutions for HSCT operations.

Douglas has established the objective that the HSCT must be able to operate from existing airports located

close to the centroid of demand. This way, the traveler will be facilitated by fast transportation from portal to

portal. This necessitates compatible approach speeds, touchdown speeds, field lengths, and noise characteristics. The portions of the airport infrastructure that must accommodate the HSCT consist of the airfield, terminal

area, and fueling facilities.

Three airfield characteristics will be affected: maneuvering space, clearance areas, and pavement

strength. The overall length will present difficulties in maneuvering on existing taxiway-to-taxiway and runwayto-taxiway intersections and also in the terminal area. Figure 37 illustrates a characteristic comparison of the

wide-body jets, the advanced subsonic transport, the Mach 3.2 and Mach 5.0 concepts. Both HSCTs present

servicing difficulties due to the wing shape with the Mach 3.2 concept further disadvantaged due to door sill

heights of 25 feet to 27 feet, which are significantly higher than current aircraft. Both the Mach 3.2 and Mach 5.0

concepts will require new fuel facilities, although LNG facilities are more complex and site demanding.

ADVANCED

TECHNOLOGY MACH 3.2

M D-11 747-400 S U BSONIC (D3.2-3A) MACH 5.0

FUEL jp Jp JP TSJF METHANE

SPAN (FT) 169.5 211.0 195.0 121.2 136.7

LENGTH (FT) 200.9 231.8 189.6 315.0 297.7

HEIGHT (FT) 57.8 60.2 62.5 60.5 49.0

TOGW (LB) 602.5 850.0 397.0 769.0 1,213.0

FIGURE 37. AIRCRAFT CHARACTERISTICS

Pilots maneuver their aircraft by maintaining the cockpit over the taxiway centerline at most ICAO airports. In the U.S., pilots maneuver by attempting to reduce the main gear track radius. If the nose gear is located

significantly aft of the cockpit, pilots may prefer to use the cockpit-over-the-centerline technique. Use of an

ottboard closed-circuit television (CCTV) may solve the guidance visibility problems and by maneuvering the

Mach 3.2 aircraft with the nose gear over the centerline, the fillet requirements may be no more extensive than

that required for stretched versions of the MD-11. Worldwide category IIIC airport capability, together with

nonvisual guidance to gates could result in nose gear-over-the-centerline maneuvering during zero-zero visibility conditions.

The only airfield clearance problem that may exist concerns close parallel runways (700 feet center-tocenter), such as at Los Angeles. Under ICAO rules, the taxi holding line is 274.3 feet from the runway centerline,

thus a 315-foot-long (Mach 3.2) aircraft will not be able to hold between the two runways without restricting

operations.

The goal to not exceed loads of current aircraft determines the number of tires and their spacing of HSCT

vehicle concepts. Further, the Mach 5.0 airplane will be restricted from current bridges and overhangs that have

not been designed for aircraft weighing more than one million pounds.

TSJF used for the Mach 3.2 aircraft is assumed to involve special handling to control contamination based

on engineering specifications regarding JP-7 fuel facilities. Therefore, new storage, distribution, and dispensing

facilities will be required for the exclusive use of TSJE Fueling and handling spills and leaks is similar to Jet A.

LNG, (Mach 5.0 aircraft) will also require new storage, distribution, and dispensing facilities as well as an on-air35
./

port liquefaction facility requiring a large, dedicated ground area and a heat-exchange tower. LNG fueling procedures will differ from current by requiring the addition of a boiloff return line from the aircraft. Grounding

the aircraft to prevent spark generation will be similar to subsonic aircraft.

Special training for TSJF deoxygenation or inerting, or handling LNG including leaks or spills will be

required. Ground crew will use gloves for protection against cold LNG temperatures (-259°F).

Fluor Engineers and Air Products performed a study of airport fuel facility requirements comparing the

costs of TSJF with LNG. Capital and operating costs of the fuel facility system(s) were developed and included

in aircraft worth determinations.

Terminal gate facilities will undergo changes to accommodate the HSCT due to fuselage length and height.

Most major terminal gate parking areas handle aircraft that are no longer than 231 feet with door sill heights

up to 17.6 feet. Consequently, the HSCT will have to be angle parked, as illustrated in Figure 38, at existing

gate areas modified for the very high door sill heights. As is shown in Figure 38, angle-parking does not require

additional terminal frontage beyond that required for the 747-400. All aircraft servicing will require special terminal building and service equipment due to the high elevation of doors and access panels. Current loading

bridges cannot be raised to the Mach 3.2 door sill heights because the bridge ramp would exceed the maximum

8-percent slope. HSCT passenger loading will probably take place from a third level added to the terminal/

concourse building. The delta or arrow-wing planform makes it difficult to access mid-fuselage doors for servicing. A special, very tall galley servicing truck will be required because of (1) close proximity of the engine nozzles

space and (2) the height of the fuselage.

FIGURE 38. GATE PARKING

Aircraft servicing will almost exclusively be turnaround activities rather than quicker through-stop activities. Today's international flight turnaround scheduled time for high-capacity aircraft is 2 hours. Economic analyses indicate added benefit of shortened turnaround times for the HSCT. A goal of I hour has been established.

Total time is projected at 75 minutes, utilizing automated self-test sequencing and by assuming that passengers

can be onboard during the systems self-testing, then the turnaround time can be reduced by 10 minutes, which

is close to the 1-hour goal. Discussions with the airlines have highlighted the desirability of achieving this goal.

Scheduled NAS Plan increases in automation of ATC navigation and communication equipment and procedures should provide efficient safe routing of all commercial aircraft including the HSCT. It is assumed that

the HSCT will not have any special approach or departure priorities and as such will experience ground and

approach delays equally with all other transport aircraft.

The following will enhance HSCT integration with the future ATC environment, reduce flight crew workload, and improve economic viability:

36
• Automaticseparationassurance, threat-alert,andcollision-avoidance

• FuelAdvisoryDepartures(FAD)tominimizeenginetime, fuel consumption, and airspace congestion

• Traffic flow metering and spacing

• Fuel-efficient flight through Automatic En Route ATC (AERA)

• Automated Dependent Surveillance (ADS) in lieu of flight crew reporting

Additionally, the HSCT must achieve international operational, regulatory, and political acceptability.

The HSCT could use the Global Positioning System (GPS) effectively as an en route navigational service where

current navigational aids over water are inadequate; high-altitude cruise of the HSCT will alleviate the subsonic

en route airspace. Global acceptance of an HSCT crossing international borders and flying within international

airspace must be obtained. Regulations regarding flights above the current level of positive-controlled airspace

(above 60,000 feet) must be developed.

Expanded high-altitude global weather information will ensure safe and comfortable flight. Flight-time

reductions will result in more reliable weather predictions, thus minimizing flight delays and rerouting, and

favorably effect fuel-reserve planning.

Environment

Community Noise. The large majority of the subsonic fleet will be Stage 3 at the time the HSCT enters

service. Hence, it was assumed that the HSCT also must meet FAR Part 36, Stage 3 noise certification requirements. The airlines generally agree that Stage 3 is an appropriate HSCT design goal. The airlines also believe

that HSCT development will require international cooperation (e.g., within ICAO) to achieve an acceptable

HSCF community noise standard. The public and regulatory agencies have more than 20 years' experience with

Effective Perceived Noise Level (EPNL), and Douglas is of the opinion that EPNL should remain the certification noise metric.

Noise reduction concepts include: inverted velocity profile (4-6 EPNdB), suppressor (6-8 EPNdB), suppressor and ejector (7-15 EPNdB), thermal shield (2-4 EPNdB), and porous centerbody (2-5 EPNdB). Some of

these concepts have been verified. However, the tests were not conducted at the higher HSCTjet temperatures

and velocities and therefore, more tests are required to extend existing data bases. Combinations of these concepts could provide 10 to 20 dB noise attenuation.

Noise level estimates at the FAR Part 36 reference locations, including the effects of the inverted velocity

profile (IVP), a jet noise suppressor, and a treated ejector are shown in Figure 39. These aircraft are sized to

meet the 6,500-nautical-mile range goal as previously discussed, but fail to meet the FAR 36, Stage 3 noise regulation. The ejector may not be compatible with the Mach 5.0 concept SERN. Both HSCT concepts show the

largest exceedance of Stage 3 requirements at the sideline reference location.

The Part 36 sideline noise estimates have assumed 12-dB and 5-dB suppression for the VSCE (Mach 3.2)

and VCE (Mach 5.0) concepts, respectively. The sideline noise levels for the Mach 3.2 concept exceed the

Stage 3 requirements by 9.5 dB. The Mach 5.0 concept sideline noise is 13 dB above the requirement. An additional 2- to 3-dB reduction in sideline noise could be achieved with operational procedures where engine thrust

is reduced early in the flight path.

P&W predicted sideline noise for both an unsuppressed VSCE and with an outer stream jet noise suppressor. Variable fan and core jet areas are key features of the VSCE which enable "high flowing" the engine. The

major noise sources are the jet (mixing plus shock) noise and the duct burner combustion noise. P&W estimated

a 4-dB benefit from a circumferential Thermal Acoustic Shield. However, jet noise dominated VSCE are projected to exceed the Stage 3 sideline noise limit.

For engines dominated at takeoff powers by jet noise, one means of reducing sideline noise is to oversize

the engines (increased airflow, diameter, and thrust) and operate them at a lower relative power takeoff condition. The noise penalty associated with increased size is more than offset by the attendant reduced jet velocity.

However, the larger engines are heavier and do not operate at optimal power in the cruise regime, thus having

increased fuel burn and either an aircraft takeoff gross weight penalty or range penalty. Increasing the Mach 3.2

engine size by 110 percent reduced the sideline noise by up to 19 dB, but resulted in an aircraft range loss of

1,400 nautical miles. An aerodynamic improvement of 20 percent reduced takeoff and approach noise levels,

but did not affect sideline noise. For the Mach 5.0 configuration, decreasing the sideline noise by 9 dB resulted

in oversizing the engine by 164 percent.

GE made sideline noise estimates for the Mach 5.0 concept with a 2D-CD wedge SERN nozzle of 112 dB,

or -- 9 dB over the Stage 3 limit of 103.0 dB.

37
CONCEPT

D-3.2-3A

ENGINE

P&W VSCE

STAGE 3 REQUIREMENTS

D-5.0-15A GE VCE

STAGE 3 REQUIREMENTS

CONCORDE OLYMPUS 593

TOGW

(LB)

769,000

1,213,000

385,000

SIDELINE

112

(- 12)

102.5

116

( - 5)

TAKEOFF

(CUTBACK) APPROACH

110

(- 8)

105.4

N/A

106

(-6)

105

N/A

103 106 105

112.0 119.5 117.0

NOTE: ABOVE NOISE ESTIMATES DO NOT INCLUDE SHOCK CELL, DUCT BURNER, OR

TURBOMACHINERY NOISE

( ) SUPPRESSION ASSUMED EXCLUDING IVP

INCLUDES INVERTED VELOCITY PROFILE PLUS SUPPRESSION DEVICES

FIGURE 39. ESTIMATED FAR PART 36 NOISE LEVELS (EPNdB)

The impact the Mach 3.2 concept would have on community noise can be gauged on the 100-EPNdB

(approximately 87-dB(A)) noise contours using the Part 36 takeoff procedure in comparison to those for a typical

long-haul subsonic airplane (Figure 40). This shows the HSCT, with the IVP and suppression devices, to be considerably noisier than the 747-200; 2.2 square miles for the 747-200, compared to 8.7 square miles for the HSCT.

In order for the Mach 3.2 concept to be comparable, a further noise reduction of 8 dB and 7 dB at sideline and

takeoff, respectively, is required and should be the basis of further innovative research. The effect of a 2:1 oversizing of the unsuppressed engine results in reducing the 100-EPNdB contour area to 3.5 square miles. As stated

above, oversizing the engine results in an unacceptable aircraft range reduction (more than 1,000 miles). The

addition of a noise-suppression device would reduce the contour areas by approximately 50 percent. The total

HSCT contour area could be reduced in size by optimizing the takeoff procedure for minimum area rather than

flying the Part 36 cutback procedure.

RANGE = 6,500 N MI

_. _-:-2-._---'_----'_---..'.-'-_--_'.'_-_'-'.'-" _ . 747-200 I HSCT _

FIGURE 40. MACH 3.2 CONCEPT 100 EPNdB NOISE CONTOURS

Sonic Boom. Sonic booms caused by supersonic aircraft exhibit vastly different acoustic characteristics

than subsonic aircraft flyover noise. A sonic boom is a high-energy, impulsive sound that generates a large subaudible, low-frequency component. Sonic boom levels are typically measured by peak or maximum overpressure. However, human response to sonic booms is a function of the entire waveshape, not just the peak overpressure (References 5 and 6). Acceptance of individual sonic boom events is best measured with a loudness

descriptor. The best metric appears to be Stevens (Mark VII) model for Perceived Loudness Level in PLdB

(Reference 7) because of its sensitivity to changes in waveform shape and inclusive research.

Listeners in a variety of conditions and locations, both indoors and outdoors, may not be sensitive to the

fine changes in waveform shape to which a loudness descriptor is sensitive. Community response to multiple

daily sonic boom events is best measured with a Day-Night Average C-Weighted Sound Exposure Level (1_%,)

38
(Reference 8). This metric is based on Oklahoma City tests and verified by the Army Construction Engineering

Research Laboratory (Reference 9). Based on this work, Douglas has selected sonic boom design criteria of a

Perceived Loudness Level of 90 PLdB and a C-Weighted Sound Exposure Level of 102 dB. However, research

on human response to sonic booms at this level must be conducted to verify that these levels are acceptable.

Sonic boom waveforms for HSCT configurations were predicted using two different methods: (1) The

sonic boom for the Mach 3.2 concept was calculated using MDBOOM which generates volume and lift distributions along the Mach cutting planes. Atmospheric propagation using Thomas' method (Reference 10) yields

sonic boom waveforms on the ground. (2) A nonlinear numerical code was used for the Mach 5.0 concept in

view of the strong shocks.

The sonic boom waveforms and their corresponding response metrics are shown for each configuration

at the beginning and end of cruise in Figure 41. As can be seen, the sonic boom design criteria are exceeded

even at the end of cruise. Except for the climb out, the most critical sonic boom occurs at the beginning of cruise

directly under the flight track. As the aircraft decreases in weight and increases in altithde over the course of

a missiori, the sonic boom levels drop significantly. At Mach 3.2, the beginning to end-of-cruise level drops from

1.9 psf to 1.2 psf, and at Mach 5.0 the level drops from 2.0 psf to 1.6 psf. To put these boom levels into perspective relative to current aircraft, the Concorde sonic boom overpressure is 2.0 psf and is banned from overland

flight supersonically. The HSCT flies at Mach 3.2 relative to the Concorde at Mach 2 and carries three times

the payload at twice the range at sonic boom levels that are slightly less. The aircraft which cannot meet boom

design goals and fly supersonically overlan,.: at the beginning of cruise may be able to do so at some point along

its flight path.

HUMAN RESPONSE METRICS HUMAN RESPONSE METRICS

3.2 5.0 _ 3.2 5.0

P, PLdB 102.1 100.7 P, PLdB 95.9 97.2

LCE, dB 107.4 106.1 LCE, dB 103.0 103.8

&P MAX, PSF 1.9 2.0 _P MAX, PSF 1.2 1.6

J

LL

_o

W

nw

no_

0c

w

0

BEGINNING OF CRUISE

M = 3.2: ALT = 65,708 FT, GW = 669,000 LB

.... M = 5.0: ALT = 84,048 FT, GW = 945,291 LB

I

% I %

'_,. _ m ,,-J

2.0 _,%

%%

1.0

0

- 1.0

-2.0

END OF CRUISE

M = 3.2: ALT = 76,173 FT, GW = 354,500 LB

.... M = 5.0: ALT = 92,213 FT, GW = 640,304 LB

%

%%

I I I _ 1 1 1__

100 200 300 400 500 600 700 100 200 300 400 500 600 700

TIME (MSEC) TIME (MSEC)

FIGURE 41. SONIC BOOM WAMEFORMS

The sonic boom off-track carpets at the beginning of cruise for the Mach 3.2 and Mach 5.0 concepts are

shown in Figures 42 and 43, respectively. Off-track boom overpressures decrease significantly; more than onehalf of the boom carpet for the Mach 3.2 concept experiences an overpressure of less than 1.0 psf, even though

39
thecenterlineoverpressureis 1.9psf.TheMach5.0conceptoff-track boom overpressures decay at a slower

rate. The large width of the boom corridor (52 miles at Mach 3.2, 65 miles at Mach 5.0) indicates that the

off-track boom levels are important and need to be considered for environmental impact.

2.0

1.6

1.2

f.

_ 08

0

o-, I

0.0 [

25,0

I I I I I

15.0 50 0.0 5.0 15.0 25.0

OISTANCE OFF TRACK. MILES

FIGURE 42, MACH 3.2 CONCEPT SONIC BOOM OVERPRESSURES ON THE GROUND

2.0

1.6

1.2

0.8

0.4

0.0

FIGURE 43.

I I I I I I I

30.0 20.0 10.0 0.0 10.0 20.0 30.0

DISTANCE OFF TRACK, MILES

MACH 5.0 CONCEPT SONIC BOOM OMERPRESSURES ON THE GROUND

Sonic boom minimization can be approached from two different perspectives: (1) Design and operate the

aircraft such that the shock strength, and hence peak overpressure at the ground, is minimized. This results in

aerodynamically slender, long vehicles. (2) Sonic boom shaping which involves creating sonic boom waveshapes

with minimum shock strengths and reduced high-frequency energy. This method is considerably more difficult

from a design standpoint and will result in aircraft configurations that depart considerably from aerodynamically

optimized designs.

4O
VariousmodificationstothebaselineMach3.2conceptwereevaluated. Themostpromisingconfiguration

involvedamodificationandextensionof thewingandresultedina20-percent reductioninoverpressure. A

comparisonofthisconfigurationwiththebaseline is shown in Figure 44. Further, a 20-percent reduction in overpressure can be obtained by increasing the cruise altitude from 65,000 to 85,000 feet. A 20-percent reduction

in overpressure is available through a 30-percent weight reduction. These results indicate that a combination

of careful aircraft shaping, and high-altitude cruise weight reduction, may lead to far field N-waves of 1.0 psf

or less. The best prospects for sonic boom minimization appear to be in a low-drag, long, slender, high-altitude

aircraft. The combination of shaping, weight reduction, and increased altitude have potential for as much as

a 60-percent reduction over current levels. This corresponds to an overpressure of approximately 0.8 psf, which

is in the range of potentially acceptable levels.

P = 102.1 PLdB

Ap = 2.0 PSF

618,000-L

65,000-FT ALTITUDE _

• 5 CONCEPT

LCE = 105.6 dB

P = 99.7 PLdB

AP = 1.6 PSF

FIGURE 44. MACH 3.2 SONIC BOOM MINIMIZATION

Emissions. Early in the HSCT study, atmospheric ozone was defined as an important element of environmental compatibility. As the study progressed, ozone concentration, atmospheric models, chemistry interactions, dynamics, HSCT characteristics, and operational factors (e.g., HSCT fleet size, altitude, geographical

markets) evolved into a process for projection of long-term effects on ozone and identification of technology

drivers and goals. NASA focused on creditable atmospheric models that would be available to interface with

study activities to evaluate HSCT concepts in the process of determining technology requirements.

One of the key goals of the propulsion system studies was achieving low levels of NOx emissions while

maintaining desired levels of thrust and efficiency. Current technology engines do not offer the prospects for

enyironmentally acceptable HSCTs from the emissions standpoint. Additional technology must be developed

to reduce the magnitude of harmful combustion products.

41
BothGEandP&W,undersubcontracts, conductedstudiesoftheimpactofadvancedtechnologycombustorsonengineemissionsandengineperformance. Theengineexhaustproducts-- specifically the nitrogen

oxides -- were estimated at supersonic cruise conditions because oxides of nitrogen emissions are most critical

at high-altitude flight condition. Here, the combustor inlet temperatures are at their highest level. Projections

were made of the exhaust gas composition and incremental engine performance for progressively more

advanced technology combustor systems relative to an engine with current combustor technology levels.

The reference engine for the Mach 3.2 concept is the P&W VSCE operating on TSJF. Many of the exhaust

gas constituents are dependent only on the composition of the fuel. These include oxides of sulphur, trace

metals, water vapor and carbon dioxide. Emissions of carbon monoxide, unburned hydrocarbons, oxides of nitrogen and smoke were estimated. The P&W VSCE with a near-term technology combustor system was projected

to have NOx Emissions Index level of 39.5 -- significantly above estimates for the 1971 U.S. SST.

Advanced combustion technology studies identified a rich-burn quick-quench (RB/QQ) concept for the

main burner. In this concept all of the fuel is consumed in a combustion zone having a very rich mixture; the

tempeniture is moderate and the formation rate of oxides of nitrogen is low. The combustion products pass

through a second reaction zone in which the mixture is lean and temperatures are sufficiently high to effect

combustion, but low enough to control the formation of oxides of nitrogen. The rich-to-lean mixture transition

is accomplished in the quick-quench section between these zones. Projections indicate substantially lower NOx

emissions -- 12.1 -- than the current technology main burner. Combustor complexity factors include length

requirements, variable geometry features, fuel preparation, and the need for novel material and/or cooling concepts for the liner in the rich zone.

A more advanced, lean_ premixed, prevaporized (PM/PV) combustion system is the most aggressive technology identified. This concet;t achieves very low NOx by not only lean burning but also by avoiding any locally

rich regions. This is accomplished by prevaporizing the fuel to a homogeneous mixture for the combustion process. This requires variable geometry of air passages to control the fuel/air ratio. The fuel must be heated to about

800°F before entering the combustor. A practical combustor using prevaporized liquid fuel has never been

demonstrated.

The estimated emissions characteristics of the P&W VSCE engine with the near-term and advanced technology burners are presented in Figure 45. The total oxides of nitrogen emissions at supersonic cruise with a

PM/PV main combustor are less than one-fourth that of the engine with the current technology main burner

(emissions index of 8.7 versus 39.5). Adding a lean, premixed, prevaporized duct burner reduces the supersonic

cruise NOx emissions to 15 percent of current technology or 6.1. A combustor concept projected to produce

the lowest NOx emissions is considered to be a very high risk approach. Substantial additional development is

required to produce viable prevaporizing and premixing systems for use with liquid fuels and the variable geometry airflow systems required for stochiometry. There are also fundamental risk elements such as preignition and

flashback in the premixing passages.

Differences in engine performance with the three main burner concepts are projected to be extremely

small with combustion efficiency above 99 percent. Some minor variants in specific fuel consumption are projected for the climb phase of the mission.

Engine emissions were processed through market models leading to analysis of ozone dynamics. One representative city-pair was selected from each of the ten international traffic regions; altitude, fuel burn, and latitude profiles were determined. Engine exhaust parameters provided by the engine companies allowed conversion of the fuel burns into quantifiable products of combustion. The fuel burned in altitude levels and latitude

bands for Mach 3.2 concept are illustrated in Figure 46; correspondingly, fleet models of fuelburn and emissions

products for the Mach 5.0 concept were developed. These data serve as inputs to atmospheric chemistry models

executed by agencies under direct contract to NASA to determine the impact on the atmosphere. The results

of these analyses will provide guidance and direction for follow-on engine technology development and specific

engine emission requirements.

42
NOx

40

30 -

20 -

10 -

0

CURRENT

TECHNOLOGY

RICH-BURN -

QUICK-QUENCH

PREMIXED -

PREVAPORIZED

ADVANCED TECHNOLOGY

FIGURE 45. MACH 3.2 ENGINE MAIN BURNER EMISSIONS INDEX

ANNUAL FUEL BURN IN MILLIONS OF POUNDS

FUEL BURN

10,000

4,000

2,000

0

100

_0

30

I._TITUOE

IDEGREES)

30

FIGURE 46. MACH 3.2 ANNUAL FUEL BURN

43
_ _ _ • _i_

L

44
CONCLUSIONS

MARKET

1. Market projections for the 2000 to 2025 time period indicate sufficient passenger traffic for ranges beyond

2,000 nautical miles to support a fleet of economically viable and environmentally compatible high-speed

commercial transports.

2. The fleet size that will satisfy these traffic projections depends on utilization, productivity, and fare premium. HSCT fleet needs of 300 aircraft in the year 2000 could grow to 1,500 or more 300-seat aircraft by

2025.

3. Between 2000 and 2025, the Pacific Rim will become the major traffic region, thus leading to a designrange objective of 6,500 nautical miles.

4. For high-speed commercial service, an increase in ticket price over competing subsonic service that would

be acceptable to the public would enhance the economic viability of HSCT. However, the price sensitivity

for coach passengers is much greater than for first class. A fare increase as small as 10 to 20 percent above

the subsonic level would sharply erode the coach passenger market.

5. Economic viability places emphasis on environmentally acceptable supersonic flight over land. The constraint of no supersonic flight over land reduces potential aircraft productivity (i.e., seat-mile per year)

by 7 to 28 percent for the Mach 3.2 concept.

CRUISE SPEED

1. Aircraft productivity increases with cruise speed up to about Mach 5 to Mach 6 for market applications

ranging from 2,000 to 6,500 nautical miles. Above this point, the relative significance of cruise speed

diminishes, and productivity is virtually constant.

2. Design mission gross weights increase with cruise Mach number and, correspondingly, advanced technology requirements are greater, possibly resulting in later readiness dates.

3. LNG applications at Mach 5.0 do not result in competitive opportunities since liquid methane's energy

content falls short of requirements. The liquid hydrogen aircraft (Mach 6) are not competitive due to the

high fuel cost.

4. Economic studies of the Mach 3.2 concept suggest viability could be achieved through modest fare premiums and successful research, providing gross weight reductions and propulsive efficiency improvements

as well as operational improvements.

ENVIRONMENTAL CONSIDERATIONS

1. Advanced engine technology has been identified that offers the potential for reductions in oxides of nitrogen to 15 percent of current levels.

2. FAR 36, Stage 3 noise certification requirements for a design range of 6,500 nautical miles cannot be met

with technology projections of this study. Oversizing engines to reduce the noise causes unacceptable

losses in mission range.

3. Concepts presented in this study are estimated to be capable of meeting performance objectives --

300 passengers/6,500 nautical miles -- with slightly lower sonic boom characteristics than the Concordc

(at 100 passengers/3,200 nautical miles). However, the HSCTconcepts fall short of meeting the tentatively

selected metric goals.

4. Sonic boom acceptability criteria are necessary to fully determine conditions of environmental compliance.

U.S. ECONOMY

1. From the standpoint of the U.S. economy, a 1,000-unit HSCT program would create a peak-employment

gain of more than 4130,000 jobs. This translates into a projected $500 billion increase in GNP and represents

an improvement in the balance of trade of more than $100 billion.

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p_(_E_INTEHTIONAhL¥ _NK
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46
RECOMMENDATIONS

RESEARCH AND TECHNOLOGY NEEDS

The HSCT study was primarily an assessment of technology in terms of potential commercial value with

a continuing emphasis on narrowing the range of Mach number design options. The purpose of the study is to

assist NASA to plan follow-on research and technology activities. Early in Phase I it was concluded that current

technology was insufficient to support a production development program. Throughout the study, technology

needs were monitored as part of the HSCT conceptdefinition process.

A compilation of insights which focuses on the Mach range suited to a kerosene-fueled HSCT is presented.

The range of interest extended to Mach 3.2 using a kerosene-based fuel, which has higher thermal stability characteristics than the currently available kerosene-based jet fuel (Jet A).

Airframe technology needs are grouped in three categories: environmental, key performance technologies, and integration and supporting technologies (Figure 47). These are based on relative priorities, significance, and program logic. The time period is predicated on a year 2000 to 2010 HSCT certification (kerosenebased fuel) with configuration development commencing in 1996-97. Research directed toward solution to the

environmental concerns is of the highest priority. It must be emphasized, however, that other technologies such

as materials and propulsion systems must be worked in the near term due to affect on HSCT weight which is

a factor in environmental compatibility.

1988 PRIORITY 1

ENVIRONMENTAL

SONIC BOOM

EMISSIONS

AIRPORT NOISE

LAMINAR FLOW CONTROl

3 YEARS

1989 PRIORITY 2

KEY PERFORMANCE TECHNOLOGIES

METHODS

AERODYNAMICS

PROPULSION

MATERIALS-STRUCTURES

HUMAN FACTORS

5 TO 6 YEARS

1991 PRIORITY 3

INTEGRATION AND SUPPORTING TECHNOLOGIES

AERODYNAMICS

PROPULSION

SYSTEMS

HUMAN FACTORS

PRODUCTION

5 TO 6 YEARS

FIGURE 47. TECHNOLOGY PLAN

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4?

REFERENCES

1. NASA Report No. CR-165723, Technology Study for Advanced Supersonic Cruise Vehicles, June 1981.

2. Morris, R. E, Brewer, G. D., NASA Report No. CR158926, Hypersonic Cruise Aircraft Propulsion Integration Study, Volumes I and II, September 1979.

3. Cubbage, J. M. and Kirkham, E S., NASA Report No. TN D-6060, Investigation of Engine-Exhaust-Airframe Interference on a Cruise Vehicle at Mach 6, January 1971.

4. Darden, C. M., NASA Report No. TPo1348, Sonic-Boom Minimization with Noseq31untness Relaxation,

January 1979.

5. Zepler, E. E., Harvel, J.R.P., The Loudness of Sonic Booms and Other Impulsive Sounds, J. Sound

Vib., 2(3), pp. 249-256, 1965.

6. Johnson, D. R., Robinson, D. W., The Subjective Evaluation of Sonic Bangs, Acoustica, 18(5), 1967.

7. Stevens, S. S., Perceived Level of Noise by Mark VII and Decibels (E), JASA 51 (2) (Part 2), pp. 575-601,

1972.

8. Borsky, P. N., Community Reactions to Sonic Booms in the Oklahoma City Area, National Opinion

Research Center, AMRL-TR-65-37, 1965.

9. Schomer, P. D., Neathammer, R. D., Community Reaction to Impulsive Noise: A Final 10-Year Research

Summary, U.S. Army Construction Engineering Research Laboratory, TR N-162, 1985.

10. Thomas, C. L., Extrapolation of Sonic Boom Pressure Signatures by the Waveform Parameter Method,

NASA TN D-6832, June 1972.

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r_

49
!
Report Documentation Page

./

National _ a_l

Space Adminis_w, Jon

1. Report No.

NASA CR-42 36

4. Title and Subtitle

2. Government Accession No. 3. Recipient's Catalog No.

5. Report Date

Study of High-Speed Civil Transports - Summary

7. Author(s)

Douglas Aircraft Company

New Commercial Programs

9. Performing O_aniz_ion Name and Address

Douglas Aircraft Company

3855 Lakewood Blvd.

Long Beach, CA

90846-0001

12. Sponsoring Agency Name and Address

NASA Langley Research Center

Hampton, VA 23665-5225

August 1990

6. Performing Organization Code

8. Performing Organization Report No.

10. Work Unit No.

537-01-31-01

11. Contract or Grant No.

NASI-18378

13. Type of Report and Period Covered

Contractor Report

14. Sponsoring Agency Code

15. Supplemen_ N_es

NASA Program Manager:

Douglas Program Manager:

Douglas Contract Manager:

Final Report - Summary

Charles E. K. Morris, Jr.

Donald A. Graf

George W. Lawrence

16. Abstract

A systems study to identify the economic potential for a high-speed commercial transport considered

technology, market characteristics, airport infrastructure, and environmental issues. Market forecasts

indicate a need for HSCT service in the 2000/2010 time frame conditioned on economic viability and

environmental acceptability. Design requirements focused on a 300-passenger, 3-class service, and a

6,500-nautical-mile range, based on accelerated growth of the Pacific region. The need for compatibility with existing airports was assumed. Mach numbers between 2 and 25 were examined in conjunction with

the appropriate propulsion systems, fuels, structural materials, and thermal management systems.

Aircraft productivity was a key parameter, with aircraft worth rather than aircraft price being the

airline-oriented figure of merit. It was determined that Mach 3.2 (TSJF) would have superior characteristics to Mach 5.0 (LNG) and recommended that the next-generation high-speed commercial transport aircraft use a kerosene fuel. The sensitivity of aircraft performance and economics to environmental constraints (e.g., sonic boom, engine emissions, and airport/community noise) was identified together with

key technologies. In all, current technology is not adequate to produce viable HSCTs for the world marketplace. Technological advancements must be achieved to meet environmental requirements. (These requirements are as yet undetermined for sonic boom and engine emissions). High priority is assigned to

reducing aircraft gross weight which benefits both economics and environmental aspects. Specific technology requirements were identified and national economic benefits projected.

17. Key Wo_s (Suggested by Author(s))

High-Speed Civil Transport

Market

Environment

Aircraft Concept

Airport Infrastructure

Economics

19. Securi_ Classification (of this report)

Unclassified

NASA FORM 1626 00T86

18. Distribution S_tement

Unclassified - Unlimited

Subject Category 05

20.'Securi_ Classification(ofthis page)

Unclassified

21. No. of pages

64

For sale by the National Technical Information Service, Springfield, Virginia 22161-2171

22. Price

A04

NASA-Langley, 1990