Technology Readiness & Development Risks of the
New
Supersonic Transport Aircraft (SST)

By
GEORGE SAOUNATSOS
COPYRIGHT 1998, ASCE / GS

Manuscript of the Technical Paper published in
ASCE's 'Journal of Aerospace Engineering', July 1998

 
1.  Abstract

Realistic specifications for the new supersonic transport call for a 2.2-Mach, 270-passenger, 5,500NM-range aircraft, requiring $15-20 billion for research and development.  Non-linear aerodynamic analysis methods promise an enhanced lift-to-drag ratio of about 45% in subsonic and 25% in supersonic cruise.  Adequate sonic-boom suppression does not appear feasible.  The resized increment in weight reduction due to the elimination ofhe droop nose is estimated at 4,500kg.  Considerable progress has been made in meeting FAR-36/Stage-3 noise requirements through the use of a Mixed-Flow Turbofan with ejector or a Mid Tandem Fan variable cycle engine.  A 5-15% lower specific fuel consumption is projected, while the dominant combustion philosophies are the Lean-Premixed-Prevaporized and the Rich-burn/Quick-quench/Lean-burn proЩcesses. Research is aimed at a NOx Emission Index of 5, resulting in less than 1% annual ozone depletion. The estimated market needs of about 550 units by 2020 justify a satisfactory return on investment (12%) for only one manufacturer.  A well structured international consortium could reduce development risks, time and expenditures through technology transfer and the sharing minimization of non recurring costs.  The Entry-Into-Service date can be placed between 2010-2015.
 

2.  Introduction

The launch of any large commercial airplane program involves a high-stakes gamble, because uncertainty always exists on whether enough units will be produced and also sold to provide a positive return on investment (ROI).  With the next generation supersonic transport this uncertainty is magnified to unprecedented levels due to the increased technical challenges.  Estimates by both American and European manufacturers suggest that non-recurring costs for this program could fall in a range between 15-20 billion US$ (Bunin et al. 1994).  Technology readiness and the understanding of the risk factors involved are key elements in the successful development of the next generation SST.

In April 1994, Aerospatiale, British Aerospace (BAe) and Deutsche Aerospace (DASA) signed a memorandum of understanding, creating a joint European supersonic research program (ESRP) and expecting the aircraft to reach operational status  by 2010.  In parallel, Snecma, Rolls-Royce, MTU and Fiat have been working together for the development of a propulsion system since 1991.  Investing no more than $12 million per year (Sparaco 1997), mainly of company funded research, the program covers materials, aerodynamics, systems and engine integration for a reference configuration by 1999.  The ESRP exploratory studies are based on a 2.0-Mach, 250-seat, 5,500 nautical mile-range (NM) aircraft.  Canards are incorporated to improve low-speed performance and handling, as shown in figure 1.
 
Figure 1: Aerospatiale’s SST concept  [SOURCE: Aerospatiale]

In the Pacific region, the Japan Aircraft Development Corporation (JADC) coordinates the efforts of Mitsubishi, Kawasaki, and Fuji Heavy Industries in their weighty research on the next generation SST.  Currently JADC is making efforts to expand their activity to 20 million dollars per year, as Hiroshi Mizuno, General Manager of JADC, mentioned to the Author (Mizuno 1996).  At the same time, Alexander Pukhov, Tupolev's chief designer, maintains that through the flight tests conducted on a Tu-144 in collaboration with NASA, Russia's research and development of a second generation supersonic transport (Tu-144) is revived.  As revealed at the Paris Air Show in 1993, the Tu-244 is a 2.05-Mach, 300-passenger, 5,100NM-range aircraft powered by four 323.8kN (72,800lbf) engines and having a maximum take-off weight (MTOW) of 349,000kg (770,000lb).  Considering, however, the current political and economic instability, Russia does not seem to be a major player in the SST development although their scientific inputs cannot be ignored.

In the United States, as a first step in assessing the market and technology needs for a viable supersonic transport, contracts were awarded in 1986 to Boeing and Douglas Aircraft companies to conduct feasibility studies.  In 1990, NASA initiated an intensive High-Speed Research program (HSR), a thorough first-order assessment of the performance benefits associated with technology improvements.  The HSR  program has been since moving along with steady support of team leader Boeing and some 50 other U.S. subcontractors (Ott 1997).  Advanced technologies in the areas of aerodynamics, structures, propulsion and flight deck systems were applied to a representative 2.4-Mach vehicle concept, called reference "H".  In Autumn 1995, engineers from NASA, Boeing, McDonnell Douglas, General Electric, Pratt & Whitney and Honeywell produced a new planform, the "technology-concept airplane" or TCA.  The TCA concept also calls for a 2.4-Mach aircraft with a capacity of 300 passengers, a range of more than 5,000NM and a MTOW of under 340,000kg (750,000lb).  NASA is currently spending a quarter of its annual $1 billion aeronautics budget for high-speed research.  By year 2002, after having spent nearly $2 billion on its HSR program (Zurer 1995), NASA hopes to have developed with the US industry the required technology for a technically feasible and economically viable supersonic transport.  This future SST may resemble the characteristics depicted in figure 2, as envisioned by the Boeing company.  Although none of the characteristics of the configuration aircraft is firmly in place, it can be observed that a horizontal stabilizer is incorporated in their baseline studies along with a double delta wing.  Notable is the extended fuselage length, which is 1.5 times larger than Concorde’s, and the corresponding increased cockpit-to-main-gear distance.  According to NASA, the convergence period towards a definitive configuration is the time period between 1996 and 1998/99, when the "Authority to proceed" will be granted.
 
Figure 2: Boeing's SST configuration  [SOURCE: Boeing]

It is the Author’s belief that among the SST concepts presented above, a lower speed (Mach » 2.0) aircraft is probably a wiser choice due to its reduced complexity, development costs and the already existing "know-how". The development cost of a 2.4-Mach SST, as extrapolated from JADC’s studies (Mizuno et al. 1991), could be 30-40% higher from a 2.0-Mach aircraft, leading to a higher production and acquisition cost.  The latter would elevate the direct operating cost (DOC) of the aircraft, requiring a ticket fare increase, while the overall operational efficiency and the shorter block-time tied to the faster cruise would be marginal (Proctor et al. 1994).  At Mach 2.4 and on transoceanic routes, for example, flight time could be reduced by about 20 minutes, but the airliner’s DOC would be significantly higher (Sparaco 1997).  As research programs progress through the technology development phases, the parties involved are becoming less willing to discuss details.  Building up more confidence in their SST technology, the United States may well become the first nation to proceed with the development of a new supersonic transport, possibly supported by Japan.  The fact that Boeing chose to stay out of the Very Large Aircraft area, although Airbus is proceeding with the $8 billion A3XX project, suggests that the American HSR team could better focus and invest more now in the Supersonic Transport’s research and development.
 

3.  Design aspects

From a design point of view, the lift-to-drag (L/D) ratio is the most important aerodynamic parameter of airliners, affecting essential economic-related performance such as maximum range, payload and fuel consumption.  The primary cause of SSTs’ high specific fuel consumption (SFC) is the dramatic fall in airplane’s L/D ratio at supersonic speeds.  Concorde, for example, experiences an L/D reduction in the order of one-half that of subsonic jets (Strack 1990).  The main goal, therefore, is to increase the lift-to-drag ratio throughout the speed regime of the next generation SST.  Methods currently being evaluated include optimization schemes coupled with state-of-the-art computational fluid dynamic (CFD) solvers, as well as nonlinearized design methods.  The development of nonlinear unsteady aerodynamic analysis methods will be used to optimize wing, fuselage, nacelle and empennage geometry, reducing drag at supersonic speeds by over 10% relative to linear design methods (Wilhite et al. 1997).  The new design and analysis tools are also pointing toward optimized wing cambers, leading to improved performance in the transonic region as well as in the desired cruise conditions.  Estimates are for an increase in L/D of 0.8 (25%) at supersonic speeds and 2.0 (45%) at subsonic ones (Ozoroski et al. 1993).  Characteristically, Aerospatiale obtained an L/D increase equivalent to 1.5 tones of payload during Mach 2.0 wind tunnel experiments (Collard 1990).  This was achieved by getting a larger laminar flow region over the wing through the optimization of its camber, sweep and twist angle.  Likewise, scientists at the UK’s Defence & Research Agency (DRA) support that they have produced a wing shape with an increase in L/D ratio of 24% at supersonic and 58% at transonic conditions relative to Concorde (Hayes 1997).  High-lift leading and trailing edge wing devices can additionally be used to decrease the required thrust for take-off, climb-out and landing, reducing in parallel community noise (Wesoky et al. 1990).  These advanced high-lift concepts when combined with modern landing and takeoff procedures such as automatic flap and throttle settings, more than double the low-speed L/D ratio compared to Concorde (Darden et al. 1993, Wilhite et al. 1997).

Supersonic laminar flow control (SLFC) can also have significant aerodynamic benefits (increased L/D) by reducing skin friction drag on SST configurations.  NASA is investigating the implementation of SLFC through a series of flight tests on an F-16XL.  The objective of this project is to achieve laminar flow over 40-50% of the wing's chord in SST cruise conditions of Mach 2.4 at 60,000ft (Smith 1995).  Program officials believe that the application of laminar flow control on SST through active suction has the potential of reducing drag by 7-9%, resulting in a considerable decrease in fuel consumption (NASA 1992, Smith 1995).  Moreover, supersonic laminar flow could lower the skin’s surface temperature by about 95 oC, allowing to push up the speed of the aircraft and still use kerosene (Jet A) as a heat sink (Strack 1990).  Nevertheless, active suction, which is accomplished through millions of nearly microscopic holes laser-drilled in the wing skin, is viewed as "high risk" technology.  It has been suggested that a 4-5% of the cruise engine power might have to be absorbed in order to drive the system (NASA CR-4659 1996).  Moreover, if one puts together the weight penalty and the structure complexity the SLFC system introduces, the benefit of laminar flow may become marginal.  It has, however, the potential of becoming the enabling technology for future supersonic and subsonic aircraft if it is partially applied on aerodynamically critical areas (Smith 1995).

Another issue that is being considered is the elimination of the droop-nose configuration.  This will considerably decrease complexity, cost, as well as drag and weight penalties through the removal of mechanisms, guide rails, hinge joints and jacks for drooping and returning the nose and visor in place.  The total resized increment in drag and weight reduction for the larger next generation SST is estimated at 4,500kg (10,000lb), which translates to as many as 50 additional passengers (Swink et al. 1992).  Without a droop-nose, however, forward external visibility is negated during take-off and landing guidance, and control must rely on the use of synthetic-vision systems (SVS).  This seems to be the case for the next generation SST, providing a de facto Category-III all-weather operation capability autonomous of the landing aids of the airports.  That, would dramatically increase the dispatch reliability and on-time arrivals, enhancing operating economics, passenger satisfaction and contributing to sizable increases in cost savings across a fleet.  A cost/benefit model developed by the Douglas Aircraft Corporation, for assessing the impacts of such enhanced operating flexibility across a typical airline fleet, estimated an average annual cost saving of approximately $188 million due to the avoidance of weather related delays (Swink et al. 1992).  Additionally, there would be an added benefit of reduced fuel reserve requirements due to the implementation of inherent Category-III landing capabilities, which could be used as additional payload or simply decrease the take-off gross weight (TOGW).  One of the major US carriers, for example, has estimated that an extra 4,500kg (10,000lb) of fuel is loaded on-board on oceanic routes across the Pacific, of which 2,700kg (6,000lb) is never burned.  NASA has completed preliminary flight tests of a synthetic-vision landing system with very encouraging feedback from the pilots (Norris 1996).  The tests were conducted on subsonic transport aircraft equipped with infra-red (IR) sensors, video cameras, X-band radar, as well as a digital-database system capable of generating a 3-D ground scene.  According to NASA, an integrated synthetic-vision system is scheduled to be tested before 2001.

The new generation SST flight deck environment will also include two-way air/ground data linking, precision satellite-based navigation and automatic surveillance.  The satellite navigation will be based on a differential Global Positioning System (D-GPS) coupled with a digitized map database, providing flight profile information and drawing the airports’ runways on the display screen (Ott 1997, Wilhite et al. 1997).  The automated surveillance will use the Traffic Collision Avoidance System (TCAS) and an on board radar for active protection from aircraft not equipped with a TCAS.  These will give SST the capability of operating within a "Free-Flight" air traffic environment as envisioned by the FAA and Eurocontrol.
 

4.  The sonic-boom issue

The availability of high speed overland corridors would have a major impact on the ability of the SST to penetrate specific city-pairs markets, involving significant overland components in the aircraft routing for which there is no effective overwater re-routing option.  Hence, the obtained reductions in trip-times and operating costs would substantially boost the commercial success of the new SST.  This leads to the essential issue of overland flight, which is the "Achilles’ hill" of supersonic transports as they are prohibited from such operations due to their sonic boom loudness level.  It is well known that conventional supersonic aircraft have a sharp nose to reduce wave drag, generating a weak bow shock.  Nevertheless, strong and multiple secondary shocks are formed.  The sonic boom phenomenon is generated by the coalescence of these shock waves in the process of propagation from aircraft to the ground (Yoshida et al. 1994).  The parameters that characterize the pressure waveform (N-wave) are the peak pressure (overpressure DP), duration and rise time (Dt).  The sonic boom intensity is proportional to overpressure (DP2), so when the peak pressure decreases the sonic boom loudness reduces.  In parallel, it has been established that extension of the rise time (Dt) decreases the component of high frequency which is more severe for human sense of hearing compared to low frequency (Archer et al. 1996, Yoshida et al. 1994).  Hence, when the rise time becomes longer, for any given peak pressure, humans feel boom noise softer.  The perceived loudness also depends on atmospheric conditions, as atmospheric gradients introduce refraction of the propagating waves which can create super-booms or double-booms.

The concept of boom minimization is based on suppressing the coalescence of multiple secondary shock waves caused by the SST in supersonic flight, so that the overpressure (DP) at ground level is reduced.  This can be achieved through the manipulation of aircraft’s design characteristics, resulting in an optimized N-wave pressure signature with significant sonic boom loudness attenuation (figure 3).
 
Figure 3: N-wave properties and optimized characteristics (exaggerated for clarity)

Conventional supersonic designs generate a non-optimized N-wave signature at ground level for all flying altitudes above 40,000 feet (Wesoky et al. 1991).  This implies that reduced sonic boom loudness can only be achieved below this specific altitude, which is not efficient for the operation of SSTs from a fuel consumption standpoint.  In contrast, low boom configurations produce a non-optimized N-wave only when the aircraft's altitude exceeds 60,000 feet, suggesting a larger operational flexibility as far as the altitude is concerned.  The main feature of a typical inexpensive low-boom aircraft configuration is characterized by a blunt nose, which can decrease the overpressure by more than 50% (Yoshida et al. 1994).  Such an aircraft shape, however, generates strong bow shock waves and increases wave drag more than a conventional sharp-nosed configuration, leading to the so called "low-boom high-drag paradox". As a result, the performance of the airplane are reduced, implying shorter range at the condition of same weight and heavier weight at the condition of same range than a sharp-nosed SST aircraft.  Alternatively, advanced low-boom configurations can be achieved by a more uniform lift distribution stretched over a longer length, so that the sonic boom maintains its weaker mid-field features with lower bow shock.  In order, however, to obtain adequate boom loudness suppression, the area distribution of an aircraft must be carefully determined aerodynamically (Vachal 1990).  Such advanced low-boom designs induce a drag penalty, tending toward lightly-loaded but larger wing surfaces (Darden et al. 1993).  Studies conducted by JADC and Kawasaki Heavy Industries indicate that low boom SST aircraft configurations, designed to cruise at Mach 2.2, would have an economic advantage only if the percent of overland distance is above 20% (Yoshida et al. 1994).  In any other case the induced lower L/D ratio would have a pronounced negative effect on aircraft performance, undermining the commercial viability of the SST.  Considering the current status in the development of the required technologies, the prognosis for a practical design in this area does not appear promising at this time (Wilhite at al. 1997, Williams 1994). For this reason, parallel studies are focusing on the potential of overland operations at somewhat reduced speeds or altitude, or over remote land areas which are either completely uninhabited or very sparsely populated.  On the legal side, the sonic boom loudness level allowed overland has yet to be clearly defined.  This seems to be more of a political rather than a technical issue, which through a "favorable" definition could offer half solution to the sonic boom problem.  Meanwhile, the operational experience of British Airways with Concorde has established that for proper secondary boom control, the aircraft should maintain subsonic cruise within 105NM from the coast during the months of unfavorable upper atmospheric conditions (Macdonald 1989).  Nonetheless, further investigation is required on the potential effects on both maritime mammals and maritime shipping from a large fleet of SSTs' overwater sonic booms.
 

5.  Propulsion Technology

In antithesis to other aircraft projects where the engine development usually leads the airframe by about two-three years, the SST program has the propulsion system and the airframe coming to life at about the same time.  The fact that almost 45% of the TOGW is fuel and the propulsion system itself is about 20% of the operating empty weight (OWE), readily points out the role of the propulsion system in the economic viability of the SST.  The challenge is to have a fuel-efficient and lightweight engine, while keeping emissions acceptably low and incorporating noise suppression devices to enable the vehicle to meet current (FAR-36/Stage-3) and future airport noise requirements.  A successful SST engine should be able to comprise:
The primary requirement, however, for the propulsion system of the future supersonic transport is an improved specific fuel consumption (SFC) in comparison with Concorde's Olympus 593 engine.  A relatively small improvement of 5-10% is expected in supersonic cruise fuel consumption due to the possibility of further increase in thermal efficiency.  This is a result of the progress made in the field of materials, allowing higher compressor and combustor outlet temperatures compared to Olympus 593 (Habrard 1989).  In subsonic operations a larger reduction in specific fuel consumption of 10-15% can be achieved, allowing to cover long enough distances in subsonic conditions.  Engineers, however, are brought against a major design dilemma since the optimum aero-thermodynamic cycle at subsonic speeds is naturally different from the optimum cycle at supersonic ones, with the first requiring a high propulsive efficiency while the latter one a high specific thrust (Hirokawa et al. 1994, Hodder 1994).  In terms of performance, the engine’s nacelle and air inlet should generate minimum drag, suggesting that the engine’s frontal area has to be as small as possible.  Theoretically, such a requirement prohibits the achievement of high propulsive efficiencies in subsonic cruise, since a small inlet area leads to a smaller mass flow rate through the engine.  At the same time, the need to obtain high specific thrust in supersonic operations introduces high exhaust velocities during subsonic cruise, which are also detrimental to propulsive efficiency.  These high exhaust velocities can reach a maximum of 900m/s.  It has been well established that the exhaust jet velocity should be in the order of 400m/s and a mean temperature of 600-800K, to guarantee an acceptable noise level and meet the take-off and approach sideline-noise requirements as depicted in figure 4 (Habrard 1989, Hodder 1994, Smith M.J.T. 1990).
 
Figure 4:  Impact of exhaust jet velocity on noise

These conflicting design requirements are summarized as following:

The difficulty therefore exists in finding a "flow multiplier" device, which at take-off and in subsonic operation can significantly increase the air flow adjusted for operation at supersonic speeds.  Two general solutions to this design dilemma have evolved.  The first method aims to optimize the engine’s supersonic cruise and control take-off noise by incorporating a silencer such as an ejector.  This, creates the problem of designing an exhaust system which gives significant noise attenuation, but with minimal thrust and weight penalty.  The best noise-suppresser devices for high speed exhaust jets developed up to now, lead to 1% loss of thrust per 2dB gained and increase the exhaust system weight and complexity (Hodder 1994).  It constitutes, however, a better solution than simply oversizing the engine and its corresponding airflow by an estimated 12%, which results to a 5% increase in the take-off gross weight (Wesoky et al. 1990, Wesoky et al. 1991).  The second solution is to use variable geometry components such as compressors with variable inlet guide vanes, auxiliary intakes which can be opened and closed as required, or variable geometry mixers and turbines.  All these can be scheduled to give low and high specific thrust cycles in order to meet performance and noise requirements.  They introduce, however, the challenge of designing novel complex components with good aerodynamic performance.  The preferred concept is likely to be the one that strikes the best balance between simplicity of design and good fuel economy.
Two candidate engines appear to have been selected by manufacturers for further assessment.  These are the Mid-Tandem Fan and the Mixed-Flow Turbofan with ejector (figure 5a,b).  Both concepts appear suitable for 2.4-Mach operations, but more work is required before either one can be finally chosen.  Snecma and Rolls-Royce have been working together on the successor of Concorde's Olympus engine since 1989, while MTU and Fiat joined in 1991.  Their cooperation has led to the Mid-Tandem Fan (MTF), a variable cycle engine that benefits from the Snecma MCV-99 and Rolls-Royce’s tandem fan concept.  The engine is equipped with a mid-bypass fan coupled to the secondary body, which allows it to operate with one or two flow paths.
 
Figure 5a:  Mid Tandem Fan concept
 
Figure 5b:  Turbofan with ejector

For subsonic conditions, the bypass flow enters the engine through two flow paths via both frontal and lateral air inlets, having a bypass ratio of greater than 2 (Lowrie et al. 1990, Proctor et al. 1994).  This double flow path, which provides low specific fuel consumption and moderate ejection speeds of about 400m/s, allows noise regulations to be met.  For supersonic cruise, the lateral inlets are closed and the variable-pitch guide vanes of the mid-fan reduce the frontal airflow into the bypass duct.  This, decreases the bypass ratio to 0.7 which is favorable for supersonic conditions (Proctor et al. 1994).  Jet speed at the nozzle outlet is then about 650m/s.  The innovation in this design is the insertion of the fan downstream of the low-pressure compressor, where the primary air flow is already compressed in a duct of a moderate, from a drag standpoint, diameter of about 2 meters (Poisson 1994).  The mid-fan therefore, plays the dual role of supplying the bypass flow in its external section, and also compressing the air in the primary circuit (between the low and high pressure compressors) in its inner section.  The Mid-Tandem Fan has a marginally better mission performance than the Mixed-Flow Turbofan, but it has a more complex design and a slightly higher weight than the latter one (Hodder 1994).  On the other hand, the Mixed-Flow Turbofan is a simpler and easier system to develop, but its main design difficulty is the long mixer-ejector nozzle which may generate installation problems on the SST.  This type of nozzle is essential as it entrains outside freestream air that is mixed with the core jet exhaust, resulting in a slower, cooler exhaust jet that reduces noise by about 16 dB (Wilhite et al. 1997).

Rolls-Royce aims to complete its studies on propulsion by year 2001, as A. Newby, Chief of Advanced Propulsion Systems at Rolls-Royce, mentioned to the Author (Newby 1996).  On the American side, a full-scale technology demonstrator engine, incorporating the materials needed for a production engine, is expected between 2001-2006.  As now defined, the baseline design for the SST propulsion demonstrator relies on a Mixed-Flow Turbofan rather than a variable cycle engine (Kandebo 1997, Norris 1996).  Meanwhile, if the current rate of the US-research expenditure is maintained, it offers a good chance of success with an attainable entry-into-service year of 2007, as A. Newby asserts (1996).
 

6.  Atmospheric Emissions

In addition to the technical challenges encountered in the development of a new propulsion system, there are several critical environmental issues involved.  While the sonic boom problem is airframe driven and the excessive airport noise levels are due to the high takeoff exhaust velocities, engine exhaust gas emissions is another major problem requiring special attention.  Important emissions from SSTs include water vapor, carbon dioxide (CO2), nitrogen oxides (NOx) sulfur oxides (SOx), carbon monoxide (CO), hydrocarbons (HCs) and soot  (NASA 1995).  These emissions are directly involved into the radiative balance of the atmosphere, making climate changes possible.  Furthermore, SSTs are expected to cruise at high altitudes of 55,000-70,000 feet, which correspond closely to the maximum ozone density.  Stratospheric ozone change will therefore take place, since SST emissions will directly affect photochemical processes which control the ozone abundance.  It has been established that photochemical loss of ozone is dominated by catalytic reactions involving Nitrogen oxides (NOx), hydrogen oxides (HOx), halogen radicals and bromine oxides.  Nitrogen oxides catalytically destroy stratospheric ozone, but also inhibit ozone destruction by HOx and halogen radical catalytic cycles through the formation of more stable gases (NASA 1995).  The effect, therefore, of adding NOx to the atmosphere is sensitive to the balance of these two effects.  Recent studies conducted by both NASA (1995) and the German Institute of Atmospheric Physics (Schumann 1994) verified that heterogeneous reactions on stratospheric aerosol particles also play an important role by reducing the ozone loss due to NOx and increasing that due to HOx and halogens.  In parallel, atmospheric circulation will affect the distribution of exhaust gases emitted by a fleet of supersonic vehicles, increasing the time spent in regions of photochemical loss, or heterogeneous reactions, and possibly enhancing their detrimental effect.

From a technical standpoint, these critical emission levels depend on the engines' injection and combustion system performance, i.e. the quality of the air/fuel mixture, its distribution in the combustor primary zone and the corresponding temperatures.  The reduction of NOx emissions remains of primary concern, having the largest Emission Index (grams of emission produced per kilogram of fuel consumed) among the substances responsible for the ozone layer depletion.  The high level of NOx emissions from current combustors is due to burning fuel near stoichiometric air-to-fuel ratios (Wilhite et al. 1997).  Hence, the key in reducing NOx production is to burn either fuel-rich or lean (figure 6).  Two combustion processes are mainly being examined; the two-stage Rich-burn/Quick-quench/Lean-burn (RQL) and the Lean Premixed Prevaporized (LPP) concept.
 
Figure 6:  The LPP & RQL combustion philosophies

In the RQL system, combustion takes place in three distinct zones, the fuel-rich the rapid quench and the fuel-lean sections.  Initially, the fuel is injected in the fuel-rich zone with a low rate of formation of oxides due to insufficient oxygen (Smith M.G. 1990).  The reaction is completed in the lean section, after being mixed rapidly with large quantities of air through the rapid quench zone.  In the lean section the temperature is sufficiently high to carry out the reactions, while avoiding higher levels at which formation of nitrogen oxides can be accelerated.  Thus, the RQL approach operates at higher fuel/air ratios than a normal stoichiometric combustion, inhibiting the NOx formation process due to the lack of available oxygen.  The alternative LPP approach prevaporizes the fuel and injects it into the air in a premixing passage, delivering a uniform droplet-free mixture to the combustion zone.  The fuel/air ratio is set as low as possible, but above stability or inefficiency thresholds.  For both combustors, liner material is a challenge for the 1,900 oC (3,500F) environment, as active cooling (with air) changes the mixing and chemistry critical for low NOx (Wilhite et al. 1997).  A mighty advantage of the LPP method is the ability of the liner to operate with small amounts of cooling air, giving an additional design flexibility (Stephens et al. 1993).  Although extensive full-scale experiments of these two concepts have not been concluded, preliminary measurements by NASA and its industry partners indicate a reduced NOx Emission Index (EINOx) of 5 (Zurer 1995).  Pratt & Whitney is focusing on the RQL concept which is capable of producing an eight-fold decrease in NOx levels compared to current engines (Proctor et al. 1994), while General Electric (GE) is examining LPP technologies.  The final selection of a combustor concept is scheduled by the American team for 1998 (Kandebo 1997).  Meanwhile, a successful LPP combustor developed by Rolls-Royce has demonstrated ultra-low NOx emission levels at one-tenth the current conventional subsonic technology (Singh 1996).  The corresponding experiments were carried out at UK’s Cranfield University, under a $1.68 million Rolls-Royce research project.  In simulated real combustion conditions the results indicated extremely low NOx levels, close to an EI of 1, as Professor R. Singh of Cranfield University mentioned to the Author.  Nevertheless, transforming the combustor success to date into fully operational engines, remains a difficult challenge in the development of the new SST.  It is yet to be shown that the above mentioned combustor technologies can go into full-scale, operating across the entire range of the speed required by a supersonic transport and still get low NOx emissions.

In an attempt to evaluate the amount of engine emissions from a fleet of SSTs, several scenarios have been examined.  Using a gas chemistry model which included the effects of both homogenous and heterogeneous reactions, the German Institute of Atmospheric Physics concluded that the global percentage change in ozone column by year 2015 is insignificant (Schumann 1994).  Similar results obtained by NASA indicate that the ozone decrease is less than 1% for the ground rules depicted in figure 7 and a NOx Emission Index of 5 (NASA 1995, NASA CR- 4656 1996, Poisson 1994, Zurer 1995).
 
Figure 7:  Annual percentage of estimated Ozone depletion

Most of these calculations, however, were performed with 2-D models which do not cover the 3-D emission distribution, the stronger variability of temperature and humidity in a 3-D atmosphere and the details of 3-D atmospheric dynamics.  Moreover, these models do not fully account for future climate changes due to the increase of greenhouse gases like CO2 or methane, implying that uncertainties are involved.  Hence, one cannot yet positively conclude that a fleet of supersonic transport aircraft will have insignificant effects on the stratospheric ozone.

Another issue under investigation is the high-altitude radiation exposure which may be of galactic or solar origin.  Galactic cosmic rays are heavy, high-energy ions that penetrate deep into the Earth’s atmosphere, unlike solar cosmic rays which are less penetrating, but can be very intense for short period of times as they are produced by solar flares.  Biological damage may therefore be imposed depending on the radiation dosage.  At supersonic cruise altitudes, the dose is double that of a subsonic airliner flying at 40,000ft., but because the trip time is halved the total radiation dosage for passengers is roughly the same.  The main concern, therefore, is for the exposure levels of flight crew members, which can be managed only partially by crew rotation scheduling based on validated radiation prediction methods (Wilhite et al. 1997).  Uncertainties also exist in our knowledge of the radiation in the upper atmosphere and various latitudes, thus research is being undertaken in an attempt to map the solar cycle effects and the maximum radiation environment.
 

7.  Structural Requirements

Optimum balance between structural performance and cost must be attained in order for the aircraft to be profitable for both manufacturers and airlines. Weight effectiveness is of prime importance in designing SST structures due to the extremely high influence of structural mass on the maximum take-off weight (MTOW). Critical issues arise in view of the extremely severe operating conditions, summarized as:
Such need for extended service-life of the future SST is outlined in the requirement for a total of 25,000 flight cycles compared to Concorde’s 7,500 estimated ones (Ermanni 1994).  Many lightweight, high temperature materials are under development including advanced aluminum and titanium alloys, polymer matrix composites (PMC), metal/intermetallic matrix composites (MMC/IMC), high temperature adhesives and high-temperature sealant. DASA believes that the Intermediate Modulus Carbon fibre may represent a satisfactory compromise between performance and costs, while toughened Epoxy-resins, BMI (Bismaleinimides) and thermoplastics have been selected for screening tests (Ermanni 1994).  In parallel, under NASA's HSR program, analytical methods are being advanced for optimizing structural designs such as sandwich, honeycomb, and superplastic-formed/diffusion bonded concepts (Stephens et al. 1993).  Polyimide carbon fiber matrix composites are also being developed to lighten the fuselage, outboard wing strake and empennage (Wilhite et al. 1997).
Engineers, however, are mainly confronted with increasing difficulties arising from high engine temperatures.  The structural total-life requirements for the SST propulsion system are the same 30,000 hours as in subsonic commercial engines.  Nevertheless, subsonic aircraft engines spend less than 10% of their mission time at the most severe engine conditions.  On the contrary, SST engines will spend about 60% of the mission time under the most severe combination of component stress levels and high temperature conditions.  The challenge is to utilize advanced materials to cope with the high temperatures without incurring excessive weight and cost penalties.  This requires significant improvements in turbomachinery, combustion, and exhaust nozzle materials.  Turbomachinery materials enhancements will involve increased life at temperature and stresses generally found in today’s subsonic aircraft engines.  The use of Titanium-based MMC and nickel-based superalloys are under consideration for use on compressor and turbine applications.  Researchers believe that they are well in track in delivering suitable materials for SST turbine blades and disks, that will probably be twice the size of the ones used in current subsonic transports (Kandebo 1997).  On the other hand, viable combustors concepts, having a long life goal of 18,000 hours, depend on the development and demonstration of a new class of high temperature ceramic matrix composites (CMC) for which no previous commercial practice exists.  A silicon carbide-silicon carbide (SiC-SiC) ceramic matrix composite seems to be the leading contender for use in the combustor, due to its excellent conductivity and thermal stress characteristics (Kandebo 1997). At the same time, light weight, high strength and high stiffness metallic, intermetallic and ceramic composite materials are being examined for the exhaust nozzle design in order to meet engine noise and weight requirements.  These include gamma titanium aluminides and thin wall castings of superalloys. In general, large complex configurations must be manufactured economically and demonstrate long life under adverse operating conditions.  However, such materials' behavior and related manufacturing processes are beyond current operating experience and only little information is available (Barbaux 1994, Stephens et al. 1993). Through intensive testing and analysis, scientists are attempting to establish their long-term performance under SST conditions, as well as their reliability and economic performance.  Although cost prediction is a complex issue due to the difficulty in providing enough data at this stage, preliminary studies from DASA and NASA indicate that SST structures can be produced at an acceptable cost-level (Ermanni 1994, NASA 1992).

Maintenance costs are an integral part of the overall commercial viability analysis of Supersonic Transports.  The labor-hours required per flight for SST airframe and engine maintenance, as experienced in Concorde, are almost 5-6 times more than in subsonic aircraft (Douglas Corp. 1989).  Nonetheless, supersonic aircraft achieve twice the productivity of subsonic jets resulting in a much lower labor cost per seat-mile.  A promising message for further reduction comes from British Airways, which has already achieved progress in Concorde's maintenance approach.  Increased component reliability allowed them to escalate service checks at 150 hours from 75, and intermediate checks at 1100 hours from 750.  Additionally, major checks are conducted every 12,000 hours with a down-time of three months (Macdonald 1989).  This notable improvement on maintenance time indicates that a significant cutback in labor-hours on the future SST is feasible.
 

8.  Development Challenges

An aircraft becomes economically viable only when the market size justifies the investment and risks that both manufacturers and airlines undertake when deciding to develop or purchase it. Apparently, the level of the investment exposure for the development of a supersonic transport is substantially larger than for a subsonic aircraft.  The development time of the SST will be longer, the amount of negative cash flow will be about 2-3 times bigger than in subsonic aircraft and a break-even point will be reached later, as illustrated in figure 8 (Bunin et al. 1994).  Thus, the project becomes vulnerable to time-dependent parameters such as market size and needs, the impact of political economic and ecological developments, fuel price, and the securing of uninterrupted financing sources.
 
Figure 8:  Investment Exposure  [SOURCE: Aerospatiale, DASA, BAe, Boeing, Douglas]

The dominant criterion, therefore, for the commercial viability of the SST is the completion of the cost-price-market loop, which translates into having the aircraft at the right time and offering it at the right price.
Closing the cost-price-market loop undoubtedly results in the designing and building of a product that is actually commercially useful, rather than merely proving its technological feasibility and superiority.  The Concorde was a characteristic example of an aircraft that did not follow this principle.  While it certainly proved the technical feasibility of high speed commercial flight, it was not affordable by the general public or the airlines.  Furthermore, it was the tripling of fuel prices during 1973 and the exciting developments in subsonic aircraft economics, best represented by the low cost per seat-mile of the wide-bodied jets, that gave the coupe de grace to the first supersonic airliner.
Competition may become a threat in the development of the future SST, guided from both direct or indirect sources.  Direct competition could arise from the possibility of two or more parallel SST programs being developed.  On the other hand, indirect competition involves the improvements in telecommunications, such as teleconferences or the Internet, and the Very Large Aircraft concept (VLA) with a capacity of more than 500 passengers and great economic advantages over long-haul routes (Donoghue 1997).  The remarkable advancements in electronic communications are more likely to affect only general business travel, which may decline by about 5% as estimated by the Boeing company (1995).  As far as the Very Large Aircraft concept is concerned, it seems that the coexistence of both large subsonic and supersonic aircraft over long-haul intercontinental traffic is plausible, as these two aircraft families will correspond to different passenger needs.  Nonetheless, the superior operating economics of the VLA, as outlined in the 15-20% reduction in DOCs from B747-400 levels (Moxon 1997) and its advantage of a much larger range relative to SST, should not be underrated.  The Boeing company, however, anticipates a potential market for just 500 VLAs over the next 20 years (Doyle 1996).  Surprisingly, the average projected market needs for SSTs in 20 years from now call for about 550 units (Boeing 1989, Fischer 1994, Mizuno 1994, Mizuno et al. 1991, Nuesser et al. 1994), closely resembling Boeing’s estimated number of VLAs required.  The amount of 550 SSTs also corresponds to the maximum achievable rate of production within a ten year period, as determined by Boeing, considering 2010 as the entry-into-service year (NASA CR-4719 1996).  This number of units justifies a satisfactory return on investment (12%) for only one manufacturer (Boeing 1989, Bunin et al. 1994), indicating no room for direct competition.

Today's global political and economic stability promises very different results from the ones obtained in the past.  The pace of change in the dynamics of the aviation marketplace is remarkable with new products, new management philosophies, shifting customer expectations and mergers that surpass national boundaries. The liberalization of the airline industry and the passenger’s demand for convenience have been the driving factors for the recently experienced market fragmentation (Dennis 1997).  New ultra long-range airliners (Boeing 777-300/200X, Airbus 340-500/600) are being developed to satisfy the increased need for convenient nonstop long-haul service, making possible new city-pair routes and markets.  The world is becoming a "global village", generating the need for rapid transit more than ever.  At the same time a potential global cooperation in the supersonic transport project appears more realistic, reflecting the political, economic and cultural unification of the world.  As Daniel Goldin (1993) of NASA contends, now is the time to make the dream of a commercially successful SST reality, as the refusal of trying to do so would be equivalent to a de facto failure.
 

9.  The case for global cooperation

The likelihood of a globally cooperative effort toward the development of a next generation supersonic commercial transport is entirely dependent on the benefits realized by the participants in such a venture.  These are identified as:
The participant countries would also benefit from the jobs created by the program as well as from the increased commerce and trade that would result by a cooperative effort and the subsequent use of supersonic transports.  A study made by the society of Japanese Aerospace Companies, estimated the potential economic benefits for Japan, in the manufacturing, trade and tourism related industries, at $30 billion by year 2020 (Iwaki et al. 1994).
Considerable knowledge has been gained on organizational and administrational needs of multinational consortiums to date.  Europe's long experience in cooperating on large projects offers good lessons about what should be done or avoided.  Concorde's, for instance, "two-headed" project organization and unnecessary duplication resulted in extra costs and delays that had a negative effect on the program's outcome (Poisson 1994).  The acquired experience, however, led to the successful Airbus Industries, Panavia and the Euro-Fighter (EFA) project, which assisted towards the modernization of Europe's main aeronautical research centers and in the production of excellent flying machines.  Transcontinental collaboration has also been proved prosperous as visualized in the CFM International, the successful US-French engine manufacturer formed by General Electric and Snecma, which has established an average 60% market share on single-aisle aircraft engines’ orders over the last few years.

At the same time, an international consortium would have to overcome several barriers such as the efficiency of management structure for the joint venture, the decision making process, the engineering design and development approach and the location of the final assembly line.  In parallel, the form of the legal entity, the process for establishing shares in the program, the allocation of work packages, as well as the sales and support mechanism add to the above.  Exchange rates, antitrust constraints, and government bureaucracy may also put restrictions in technology transfer, supplementing the obstacles that an international cooperation will probably face.  Thus, the key element for success in an international consortium should be a flexible management structure through an effective business set-up, which may take the form of a "Limited Liability company" or a "Lead company with subcontractors".  The five largest aircraft manufacturers (McDonnell Douglas, Boeing, Aerospatiale, British Aerospace and Deutsche Aerospace), through a common presentation during the 7th European Aerospace Conference, agreed that if an international program was put forth, it would follow the phasing depicted below (Bunin et al. 1994).
 
PHASE
DURATION
COST
I.     Evaluation & Planning
1990-1998
$10-12 Million
II.    Preliminary Design & 
       Development
1998-2002
$1.5-3 Billion
III.   Development & Production
7 Years to Certification
about $15 Billion
Table 1 Potential program phasing and costs

At present, activities are focused on the review of common interest and on coordinated studies including market, technical and business aspects of the project (Swadling 1992).  In parallel, it is realized that competition among design offices in preliminary project phases is vital, to generate new ideas and sharpen the creativity which is indispensable in the successful completion of ambitious projects.  Hence, independent research and development of critical SST technologies is taking place, to determine if the tough design and mission targets of the SST can be materialized.  After the bulk of this "risk reduction" work is completed, the aerospace industry will have a reasonably clear view on whether the project is feasible from a technical and environmental standpoint.  In parallel, they will be able to accurately evaluate the aircraft's commercial potential, so that the program can be launched by year 2002.  Through the potential development and production stages presented on Table 1, the first flight could be made possible as early as year 2007.  Accounting for potential delays associated with the development of such a demanding project, a more realistic Entry-Into-Service date could be placed between 2010-2015.

 

10.  Conclusions

With the anticipated technology availability of the next decade, the following SST specifications could be supported:
 

PROJECTED
REALISTIC
Average Speed
2.4 Mach
2.2 Mach
Capacity
250-300 Pax
about 270 Pax
Range
4,500 - 6,000 NM
about 5,500 NM
Engine Thrust
>223kN (50,000lbf)
- - - // - - -
SFC (Subsonic) 
SFC (Supersonic)
Reduced by 10-15% 
Reduced by 5-10%
 - - - // - - -
Aircraft Length
89-96 meters
 - - - // - - -
Aircraft Span
41-43 meters
- - - // - - -
MTOW
about 340,000kg (750,000lbs)
- - - // - - -
NOx EI
 < 5
5
Noise Level
Far 36-Stage III
- - - // - - -
 Table 2:  SST’s projected specifications
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