Supersonic transport aircraft are not new. In the 1960s and early 1970s, there were three major supersonic transport projects in Europe, the United States, and the Soviet Union. Only one of those projects, the Concorde, was completed and continues in revenue service-and only between Western Europe (London and Paris) and the eastern U.S. seaboard (New York and, until recently, Washington). Only 13 of these aircraft are in airline service, including charter operations and demonstration flights. The ground and flight operational acceptability for this type of aircraft has been demonstrated worldwide.
The life development of Concorde aircraft now suggests they could continue in service for an additional 10-15 years. The levels of exhaust emissions produced by Concorde's Olympus engines [EI(NOx) = 18] is higher than could be achieved now but is commensurate with the design knowledge of the 1960s. With a limited small fleet and an average aircraft utilization of 600 hours per year, these aircraft do not appear to constitute a major environmental concern, although they must be considered in aircraft fleet mix scenarios in view of the Concorde's higher cruising altitude.
The Concorde operates for passengers prepared to pay high fares. Improvements in materials, structural, and systems technology that are available or currently being developed could make a second generation of supersonic aircraft more widely affordable. Studies have concluded that an aircraft with between 250 and 300 seats cruising at Mach 2 to 2.4 (at altitudes between 16 and 20 km) is most likely to be successful (Shaw et al., 1997). To make such aircraft effective for the long overseas routes that benefit most from the increased speed and maintain viability with regard to viewpoint of sonic booms, the projected range must be at least 8000 km (and possibly 10400 km). Studies have examined a wide range of speeds and concluded that speeds higher than Mach 2.4 offer little gain in block time, whereas they exacerbate airframe materials and propulsion problems, hence increase technical risk (Zurer, 1995). Prior projections concluded that aircraft with a cruise speed of Mach 2.0 to 2.4 were feasible for entry into service in 2005, and a hypersonic vehicle cruising at Mach 5 might enter service by about 2030 (Zurer, 1995). Events have shown that these projections were optimistic, and it is unlikely that a new Mach 2 to 2.4 vehicle will enter service much before 2015. By the same token, required research for the hypersonic vehicle and its economics would make entry into service of a hypersonic vehicle unlikely before 2050-and possibly later unless scheduling and airport curfews could be accommodated to demonstrate higher cruise speed benefits. Therefore, the focus of the remaining discussion is on vehicles cruising at speeds up to Mach 2.4.
The Concorde has already demonstrated the practicality of Mach 2.05 as an achievable cruise speed with aluminum alloys for the basic structure. For speeds above Mach 2.2, more exotic materials would be required including titanium alloys and organic composites for structural items and more complex air intakes. At speeds between Mach 2 and 2.4, airframe characteristics currently dictate cruise altitudes between 16 and 20 km. Optimization studies are planned to investigate lower cruise altitudes, recognizing the potential benefit of minimized ozone impact. To enable the inclusion of route segments over populated areas without sonic booms, an advanced supersonic airliner must also be capable of cruising efficiently in an environmentally acceptable manner at subsonic speeds and lower cruise altitudes.
Figure 7-43: Military aircraft inventory (1992 and
The characteristics of potential second-generation supersonic transports and their consequent impact on the atmosphere differ substantially from those of subsonic transports. First among these differences is the cruise altitude in the stratosphere, which is near where ozone concentration peaks. For a given level of emissions, the supersonic transport's potential impact on the ozone column is larger than that of subsonic aircraft from an NOx-ozone depletion perspective. This impact led to strong opposition to U.S. supersonic transport development in the 1970s and, together with the potential airport noise impact and economic considerations, led ultimately to cancellation of the development. There is now consensus that emissions from a second-generation supersonic transport must be limited to levels that will have a "negligible effect" on ozone.
There is also concern regarding the effects of carbon and sulfate-based particulates, and water vapor (Albritton et al., 1993; Stolarski et al., 1995). This subject is discussed in Chapters 2 and 3. Atmospheric modeling has indicated that the effects of NOx are likely to be small for a fleet of 500 to 1,000 Mach 2-2.4 aircraft if EI(NOx) is near 5 g kg-1. This issue is discussed in detail in Chapters 4 through 6. This level of emission has become a technology development target for the second generation supersonic transport. As yet there are no such targets for particulates, but such targets may emerge before the decision time for a production development program is reached.
The overall efficiencies of subsonic and supersonic propulsion systems at comparable technology levels are not very different. The lift/drag ratio of supersonic aircraft, however, is substantially lower than that of subsonics-no more than about 9 compared to about 20 for subsonics, as is shown in Figure 7-41. This lower ratio is related to losses caused by shock waves, which can be minimized but not eliminated by sophisticated designs. Thus, for the same range the supersonic transport must carry a larger fraction of its mass as fuel. For a given weight, its engines must also produce more thrust in cruise because of the larger drag. Primarily as a result of these effects, the fuel burn per passenger mile of a supersonic transport is correspondingly large-two or more times that of a subsonic transport. Nevertheless, economic studies that incorporate time zone/ productivity considerations indicate that the greater productivity resulting from shorter block times could enable the supersonic transport to compete with subsonic transports. There is no intent to further quantify this question here. The remaining sections address the technological challenges that set supersonic transport propulsion apart from those faced by advanced subsonic engines.
This subject is covered in greater depth in Section 7.4, but the key points are reviewed here briefly with emphasis on supersonic aspects. Overall propulsion system efficiency may be regarded as the product of two factors: Thermal efficiency and propulsive efficiency. Thermal efficiency is the ratio of hot gas power produced by the engine gas generator to the power in the fuel flow. It is ideally controlled by the temperature ratio of the compression process-that is, the ratio of temperature at the discharge of the compressor to ambient temperature. In modern engines, the compressor discharge temperature is limited by the temperature tolerance of materials suitable for use in compressors, to a level of about 850 to 900 K. Ambient temperature is nearly constant at about 220 K (although the temperature varies considerably even at the altitudes flown by supersonic transports). Thus, the temperature ratio is about 4, resulting in an ideal thermal efficiency of about 3/4. The actual efficiency is lower because of various losses but remains quite high.
Propulsion efficiency, which is the ratio of power pushing the airplane to the hot gas power produced by the gas generator, ideally depends only on the ratio of jet velocity to flight velocity. It can be increased toward the limit of unity by increasing the bypass ratio. The bypass ratio that is best for any given application is determined by a balance between the increased drag and weight associated with the larger engine frontal area needed for increased airflow and the increased propulsive efficiency associated with the larger airflow. There is an additional factor in fixing the bypass ratio-the fan pressure ratio. Engines for supersonic applications require a higher fan pressure ratio than subsonics for similar improvements in efficiency. When these compromises are struck for the subsonic and supersonic propulsive systems, the propulsive efficiencies of the two systems at cruise conditions are not very different. Over the past 2 decades, the bypass ratios of subsonic engines have increased from 2 toward 10 as lighter weight and aerodynamically more sophisticated designs have evolved. For supersonic propulsion systems, the cruise bypass ratio of choice is now in the range of 0.5-1.0, whereas for the Concorde and for the U.S. supersonic transport of the 1970s it was zero.
In contrast to subsonic engines-in which the needs for low cruise fuel consumption and low take-off noise are synergistic in that they both favor high bypass ratio-supersonic engines face a severe conflict between the need for low bypass ratio in transsonic acceleration and cruise, and the need for higher bypass at take-off to limit jet noise. This conflicts leads to supersonic propulsion systems in which the effective bypass ratio can be small at cruise and relatively large at take-off. Such variation can be achieved by several different approaches, characterized at the extremes as variable turbomachinery systems and ejector nozzle systems. Common to all approaches is that the engine operates at a relatively high exhaust velocity at cruise and a low exhaust velocity at take-off.
The engine concept favored by the U.S. program is a low bypass turbofan (0.6) with a mixer-ejector nozzle for noise suppression at take-off. It is projected that a supersonic transport with this propulsion system could meet Federal Aviation Regulations 36 Stage 3 noise requirements. An alternative concept proposed by the European engine consortium has a fan at the midsection of the engine, fed by auxiliary inlets at take-off, to produce a bypass ratio of about 2 at take-off. The auxiliary inlets are closed and the fan airflow is decreased via a geometry change to give a lower bypass in cruise. This engine concept is estimated to meet the Stage 3 noise requirement with a conventional variable area nozzle such as that used on the Concorde.
For designs under current consideration, the compression pressure ratio is about 22 at take-off conditions, and the corresponding temperature rise is about 500 K, producing a compressor outlet temperature of about 800 K. By comparison, a subsonic engine at take-off would have a pressure ratio of 35 with a compressor exit temperature of 900 K. At supersonic cruise, the compressor outlet temperature is limited by the materials to about 950 K. For a subsonic cruise engine with a pressure ratio of 40, the corresponding compressor outlet temperature is about 850 K or lower, limited by practical aerodynamics. The consequence is that the supersonic engine runs hot at cruise and cool at take-off-the opposite of the subsonic engine.
At cruise, subsonic engines have an overall pressure ratio, including inlet ram pressure rise, in the range of 35-50. The overall pressure ratio of a supersonic propulsion system, including the intake pressure rise and the remaining turbomachinery pressure ratio, is in the range of 130-140, calculated as a ratio of the compressor exit total pressure divided by the inlet static pressure. Whereas the conventional subsonic system might have EI(NOx) of 10-15 g kg-1 fuel using a conventional combustor, even with the double annular system the advanced supersonic engine would barely achieve EI(NOx) of 30 (Lowrie, 1993). Combined with the cruise altitude in the high ozone concentration zone, this high NOx emission would not be acceptable for a supersonic transport. This problem is well recognized, and there are many research activities throughout the western world to develop ultra-low EI(NOx) systems. Although the combination of the 5g kg-1 EI(NOx) goal and high compressor outlet temperatures may seem formidable, research programs have shown that because of the high altitude, the low absolute pressures in the supersonic transport combustors allow easier control of the combustion mixing process. Contemporary findings therefore support the likely achievement of this target (Shaw et al., 1997).
Supersonic transports will also be subject to LTO cycle emissions rules and potential climb/cruise emissions rules in common with other air traffic. In this respect, the lower combustor inlet temperatures and pressures experienced at take-off will mean that NOx requirements in the airport neighborhood could be met more easily. The combustion system will be optimized for subsonic climb and cruise as well as supersonic cruise conditions, so attainment of low emissions still requires considerable effort.
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