8
PROPULSION

INTRODUCTION

Propulsion technology offers the greatest single contribution to the improvement of cruising economy and the environmental impact of commercial aircraft. The past three generations of gas turbine engines have incorporated increased turbine inlet temperature, increased compressor pressure ratio, increased bypass ratio, improved fan and nacelle performance, reduction of noise and emissions, and improved reliability that led to a continued dominance of the world commercial aircraft market. This pace of development in new engine technologies, together with advances in engine-airframe integration, can, with adequate support from industry and the National Aeronautics and Space Administration (NASA), be continued over the next 10 to 20 years, thus providing superior propulsion systems entering service through the 2005–2015 period.

These gains rest, of course, on continued development of new and improved materials and material-processing techniques, the clearly available advances in turbomachine technology, promising progress in combustion technology, and vastly improved utilization of computational fluid dynamics (CFD) in engine design procedures. Finally, there is the unknown impact of novel technologies, such as ''smart engines'' and magnetic bearings, that may completely change the course of engine development.

This wide range of possibilities for engine development is narrowed by the general type of commercial aircraft to which they will be applied. The specific requirements differ between the advanced subsonic transports, the high-speed civil transport (HSCT), and the short-haul class of aircraft. Although these engine classes have many features in common, each has unique criteria that have a major influence on its design.

In light of the factors mentioned above, one section of this chapter is devoted to the energy efficiency, economy, and environmental goals of the engines appropriate to each of the aircraft types. Sections are then included dealing individually with technological issues that require discussion in greater detail. The boxed material summarizes the primary recommendations that appear throughout the chapter, with specific recommendations given in



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Aeronautical Technologies for the Twenty-First Century 8 PROPULSION INTRODUCTION Propulsion technology offers the greatest single contribution to the improvement of cruising economy and the environmental impact of commercial aircraft. The past three generations of gas turbine engines have incorporated increased turbine inlet temperature, increased compressor pressure ratio, increased bypass ratio, improved fan and nacelle performance, reduction of noise and emissions, and improved reliability that led to a continued dominance of the world commercial aircraft market. This pace of development in new engine technologies, together with advances in engine-airframe integration, can, with adequate support from industry and the National Aeronautics and Space Administration (NASA), be continued over the next 10 to 20 years, thus providing superior propulsion systems entering service through the 2005–2015 period. These gains rest, of course, on continued development of new and improved materials and material-processing techniques, the clearly available advances in turbomachine technology, promising progress in combustion technology, and vastly improved utilization of computational fluid dynamics (CFD) in engine design procedures. Finally, there is the unknown impact of novel technologies, such as ''smart engines'' and magnetic bearings, that may completely change the course of engine development. This wide range of possibilities for engine development is narrowed by the general type of commercial aircraft to which they will be applied. The specific requirements differ between the advanced subsonic transports, the high-speed civil transport (HSCT), and the short-haul class of aircraft. Although these engine classes have many features in common, each has unique criteria that have a major influence on its design. In light of the factors mentioned above, one section of this chapter is devoted to the energy efficiency, economy, and environmental goals of the engines appropriate to each of the aircraft types. Sections are then included dealing individually with technological issues that require discussion in greater detail. The boxed material summarizes the primary recommendations that appear throughout the chapter, with specific recommendations given in

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Aeronautical Technologies for the Twenty-First Century Recommendations General NASA must vastly strengthen its propulsion technology program to include: much greater emphasis on subsonic transport propulsion systems where the United States has lost its technological edge over foreign competitors; continued support for the HSCT propulsion program; and a strong, broad-based propulsion technology program to position the United States for the post-2000 short-haul markets, including a better balance between vertical takeoff and landing commuter systems and conventional systems. Specific NASA should increase its support for generic computational and experimental propulsion research. In addition, NASA must lead in the development of technical communication with industry in computational science applied to propulsion. NASA should take the initiative in setting up a joint NASA/industry program for innovative subsonic propulsion technology that is at least equivalent to the NASA/industry HSCT propulsion program. NASA must support the development of specialty materials not currently available as commodities that, if properly developed, will become common in aircraft engines. The basic research effort at the Lewis Research Center (LeRC) in low nitric oxide combustors for the HSCT has produced excellent results and the momentum should be maintained. NASA should put in place the planned Joint Technology Acquisition Program between LeRC and industry. NASA should take advantage of its unique position to mount a substantial program in active control technologies for aircraft engines. NASA's turbomachinery research program, should be strengthened to the point at which NASA can recapture its leadership role. NASA's LeRC should direct increased effort toward the enabling technologies in compressor, combustor/turbine, and control accessories for engines appropriate to short-haul aircraft. order of priority, and the benefits that are possible with adequate research and development effort in propulsion. PROPULSION FOR ADVANCED SUBSONIC AIRCRAFT Advances in propulsion system technology have been the prime source of subsonic transport performance improvements for more than 30 years and this trend has continued through

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Aeronautical Technologies for the Twenty-First Century Benefits of Research and Technology Development in Propulsion Engine Performance Reduced engine fuel consumption Reduced emissions Reduced engine noise Reduced engine weight Greater engine thrust Enhanced reliability Reduced maintenance requirements Long-life materials and components Design for maintenance Engine Design and Development Shortened development cycle Improved computational capabilities for propulsion Improved testing facilities for propulsion Technology validation the evolution of high bypass turbofans. Engine fuel consumption has improved more than 40 percent from the early turbojets; even within the high bypass class, Pratt & Whitney (P&W), Rolls Royce, and General Electric (GE) engines have improved by 16 percent since 1970. Size is a major factor in aircraft productivity and in fuel burned per seat-mile, and the growth in engine thrust has been a major contributor to aircraft size. The high bypass turbofans of 40,000 pounds force (lbf) thrust ushered in the jumbo jet era, and engine growth to 50,000–60,000 lbf thrust has led to even larger, more capable aircraft. Outstanding improvement in engine reliability has resulted in long-range (over water) twin-engine aircraft with engines of 75,000–85,000 lbf thrust under development for mid-1990s service. Growth to 90,000–100,000 lbf thrust is expected. State of the Art The state of the art in new large engines is characterized by an installed propulsion system cruise-specific fuel consumption of 0.54–0.56, including nacelle drag but not customer bleed. The installed thrust-to-weight ratio is about 4.5. Cycle pressure ratios are 36–38; the compressor discharge air temperature is 1250ºF; and the turbine rotor inlet gas temperature is 2250ºF, with the mechanical design redline 150ºF higher. Materials include fourth-generation nickel superalloys, high-strength titaniums, and carbon composites. Controls are digital. Most turbomachinery component efficiencies are 90–92 percent; compressors have greater than 87 percent polytropic efficiency. Large engine fans are 8–10 feet in diameter; fan inlets and cowls

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Aeronautical Technologies for the Twenty-First Century are 11–14 feet in diameter and 15–18 feet long. Prices for large engines are around $100 per pound of thrust, and because of intense competition, industry will make major efforts to reduce this. In the 1990s this technology can be expected to evolve to 50:1 cruise pressure ratio, with turbine inlet temperatures up 100–150ºF as further materials improvements are made. Refinements will continue to be made in component performance, nacelle performance, and weight. At this point it is worth mentioning a program that exemplifies the type of joint NASA/industry cooperation that can have significant impact on the state of the art and on the overall competitiveness of U.S. products. The Aircraft Efficiency (ACE) program was begun in the 1970s in response to the energy crisis and contained six separate programs aimed at producing real improvements in the efficiency of aircraft. One of the six ACE programs, the Energy Efficient Engine (E3) program, had a goal of 12 percent reduction in engine fuel consumption. The program ran from 1977 to 1982 and was jointly funded by NASA, GE, and P&W. The goal was met and, most importantly, both GE and P&W have aggressively incorporated the component and systems technology that resulted from the E3 program into their current generation of engines, including the GE CF6-80 and the P&W 4000 engines. Also, the GE-90 engine that will be certified in 1994 is very closely related to the engine that was developed and tested during the E3 program. The Committee believes that this type of program can provide tremendous benefits to the U.S. aeronautics industry and should be pursued wherever feasible. The Future In 1989, NASA Lewis Research Center (LeRC) sponsored initial preliminary design studies of a new class of high thermal efficiency, high propulsive efficiency engines. Pressure ratios of 75–100:1 and turbine rotor-in temperatures of 3000–3400ºF were explored, using new materials and technologies, for the post-2000 period—far beyond the evolution expected in the 1990s. From this initial work it is apparent that there could be one block of new technology readied in the next 10 years for service around 2005 and a second block, less defined at this time and more dependent on new fiber-reinforced materials, that could be available about 10 years later for service in 2015. Both would permit dramatic advances. The first (2005), would yield a 15–20 percent reduction in fuel burned by mid-1990s engines, resulting from reduction in fuel consumption and reduction of propulsion drag and weight. The aircraft direct operating cost (DOC) could be reduced by 8–10 percent (at a fuel price of $1.00 per gallon). The second (2015) could lead to 30 percent reduction in fuel burned and 12–15 percent DOC reductions from mid-1990s levels. These advances are greater than the high bypass turbofans of the late 1960s. To achieve these results, higher core engine thermal efficiencies and higher propulsive efficiencies must be combined with improved nacelle and installation technology—all at lighter weight and at affordable costs. Figures 8-1 to 8-3 show trends in pressure ratio, compressor exit temperature, and turbine inlet temperature, respectively. The factors contributing to these three items are summarized in Tables 8-1, to 8-3.

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Aeronautical Technologies for the Twenty-First Century FIGURE 8-1 Overall pressure ratio for high-bypass ratio turbofans, maximum climb. The core engine technologies that are key to the improvements listed in Table 8-2 are improved metal materials, coatings, and fabrication techniques (2005); new engineered materials such as titanium metal matrix composites (MMCs), titanium aluminides, and nickel aluminide (2005); fiber-reinforced metal and ceramic high-temperature materials (2015); more efficient turbine blade cooling; possibly cooling air coolers;

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Aeronautical Technologies for the Twenty-First Century FIGURE 8-2 Compressor discharge temperature for high-bypass ratio turbofans, sea level static. advanced, more efficient, high-speed turbomachinery; turbomachinery with "smart" controls (2015); and combustors to reduce takeoff emissions.

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Aeronautical Technologies for the Twenty-First Century FIGURE 8-3 Turbine inlet temperature for high-bypass ratio turbofans, sea level static. The corresponding keys to providing the improvements in low pressure systems as shown in Table 8-2 are

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Aeronautical Technologies for the Twenty-First Century TABLE 8-1 Factors Providing Improved Core Thermal Efficiency   2005 Service 2015 Service Higher cycle pressure ratio 60–75 >75 Higher turbine rotor-in temperature 2900–3000°F 3200–3400°F Higher component efficiencies 1–2% 2–3% More efficient customer power and bleed extraction Advanced ECS Power by wire TABLE 8-2 Factors Providing Improved Propulsion Efficiency   2005 Service 2015 Service Higher fan efficiency 2% 2–3% Higher fan turbine efficiency at higher loadings 1% 2% Increased bypass ratio Yes Yes integrated fan aeroacoustics for higher efficiency and lower source noise; advanced materials for higher-temperature, lighter-weight fan turbines; composite fan blade development; advanced gearbox systems employing new materials, bearing technology, and lubricants; simpler higher-reliability variable pitch actuation systems; and lower-cost designs and manufacturing. Current engines have bypass ratios of 5–6 at fan pressure ratios of 1.65–1.8. With the higher turbine temperatures of the post-2000 period we must explore a range from direct-drive, mixed-flow, high-bypass turbofans (1.55–1.7 fan pressure ratio, 8–10 bypass ratio) to geared-drive, very high bypass engines (1.35–1.4 fan pressure ratio, 15–20 bypass ratio) with fixed-pitch fans and variable-pitch reversing fans. The range of engines and applications should also include advanced gear and turbine-driven unducted fans (1.1 fan pressure ratio and 35–50

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Aeronautical Technologies for the Twenty-First Century TABLE 8-3 Improvements in Nacelle and Installation Technology   2005 Service 2015 Service Reduced fan cowl drag 2% More ? Eliminate interference drag Yes Yes Better acoustic treatment—less noise Yes Yes Weight reduction Yes Yes Integrated advanced ECS and laminar flow suction system Yes Yes ultrahigh bypass ratio). As bypass ratio increases and fan pressure ratio decreases, specific fuel consumption improves but fan diameter increases dramatically. Nacelle and installation drag increases and weight increases. The low-pressure fan, fan turbine system, and the composite lightweight nacelles weigh two-thirds of the entire system in today's power plants. In past studies, such trends increased cost, eroded most of the fuel efficiency benefits, and drove up the direct operating cost. The question for the future is whether new gearbox, acoustic, and nacelle technology can shift the optimum bypass ratio for better economics at lower noise. Given the state of the art of large gearboxes, it seems questionable that engines of 15–20 bypass ratio would first enter service in the 60,000–100,000 lbf thrust category or for any long-range twins over water. The same holds for unducted ultrahigh bypass ratio turbofans, which have the lowest fuel burned potential. In neither case should this discourage very careful examination of their potential for the 2005 and 2015 periods. The keys to providing these improvements in nacelle technologies are integrated nacelle, wing and fan aeroacoustic development; advanced mixers; smart controls; advanced composites and lighter-weight integrated nacelle, reverser, and fan structures; laminar flow nacelle; advanced aircraft environmental control system; and lower-cost designs and manufacturing.

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Aeronautical Technologies for the Twenty-First Century Future noise reduction is possible but must be balanced with economic improvements. Source noise reduction combined with advanced nacelle acoustic features is necessary. An aggressive target is 10 dB below Federal Aviation Regulation (FAR-36) Stage III restrictions. Observations for Advanced Subsonic Aircraft It is clear that through both innovation and evolution in the twenty-first century, propulsion will provide technological improvements to transport aircraft equal to those accruing from aerodynamics, materials and structures, and active controls combined. Very large gains can be made in performance, weight, noise, emissions, and operating costs. U.S. engine products have the greatest share of the large subsonic propulsion market—about 80 percent in 1990—but the United States no longer has a technology edge. The U.S. propulsion industry is focusing on the next 10 years—the next new products and derivatives. Manufacturers have committed billions to research and development in the 1990s and to improved manufacturing capabilities. They have few financial resources left to undertake the high-risk, high-payoff technologies for advanced post-2000 products. The subsonic transport market, which is by far the largest, the most certain, and the most competitive, receives disappointing support from NASA. There is no evidence of a coherent, comprehensive approach. Only a few selected elements of subsonic technology are under way or in planning, and these mostly in-house. There are no goals for maintaining a competitive edge. The investments made by NASA in subsonic propulsion technology through joint NASA/industry programs in the 1975–1983 period were very effective. The 1989 preliminary design studies with engine companies on high thermal efficiency and advanced materials were a good start, but small and with no follow-on. This valuable type of investment has been drastically reduced over the last eight years and, as a consequence, NASA is not playing a strong "pathfinder role" in subsonic propulsion technology. Recommendations for Advanced Subsonic Aircraft It is the belief of the Committee that the technology for subsonic transport propulsion must be vastly strengthened; there is great opportunity here. To do so, NASA should take the initiative in setting up a joint NASA/industry program for innovative subsonic propulsion technology that is the equivalent of the NASA/industry HSCT program. There is no reason to do less. The first step should be to set up a vigorous preliminary design activity for aircraft propulsion systems—again, the equivalent of what has been underway for HSCT. Systems for two time frames should be examined: year 2005 service and year 2015 service. Improvement goals for fuel burned, DOC, noise, and emissions should be set. Firm conclusions can be reached regarding benefits and the definition of key high-risk/high-payoff technologies described earlier in this section. Programs for enabling technologies can then be planned.

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Aeronautical Technologies for the Twenty-First Century One of the greatest benefits of this preliminary design effort is that it would get the aircraft industry, the propulsion industry, and NASA working together as a team to address high-risk, high-payoff technology in a disciplined system context, 15 years before products are expected. PROPULSION FOR HIGH-SPEED CIVIL TRANSPORT The HSCT's three key requirements, on which the NASA/industry program is focused, are economic viability relative to advanced subsonic transports of the same time period, FAR-36 Stage III noise standards, and insignificant depletion of stratospheric ozone. The consequent principal challenges to the propulsion system are price and fuel consumption, weight and noise, and emissions. Engine weight and noise go together because the thrust level and jet velocities are fixed by the aircraft's takeoff and supersonic cruise characteristics. The engine contenders are in the 650 pounds per second airflow, and 50,000–60,000 lbf thrust class. The HSCT will probably require four such engines, with takeoff jet velocity in the range of 2,400 to 3,000 feet per second. Four candidate engine types which are being sorted out by the preliminary design process: a turbine bypass engine—basically a single spool turbojet with a ''variable cycle bleed'' of compressor air around the turbine; this engine would use an ejector nozzle to entrain another 100 percent or so of air to mix with the engine exhaust during takeoff; additional suppression features are included; a mixed-flow turbofan similar to U.S. Air Force/U.S. Navy fighter engines, but with two-dimensional ejector nozzle of more than 60 percent extra air entrainment, mixers and suppression features; a variable-cycle engine of the general sort that GE flew in the U.S. Air Force (USAF) Advanced Tactical Fighter prototypes, but with higher bypass and a two-dimensional ejector/ mixer/suppressor nozzle; and a version of the mixed-flow or variable-cycle engine with special fan features for bringing additional air aboard. Generally, these engines have cycle pressure ratios between 20:1 and 25:1, consistent with compressor discharge temperatures of about 1250°F, which is an important consideration for long-life supersonic cruise. Takeoff turbine temperatures will be 2900–3000°F, with redline design temperatures 100–150°F higher, and supersonic cruise temperatures about 2700°F. Engine controls will be advanced digital. To meet the FAR-36 Stage III requirements it is necessary to bring aboard a large additional amount of air during takeoff to reduce high-velocity jet noise through mixing or shielding, in addition to mechanical noise suppression devices. These features add a great deal of weight. The total gross weight penalty charged to propulsion noise reduction is 50,000–100,000 pounds with current concepts. The larger fuel consumption associated with the

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Aeronautical Technologies for the Twenty-First Century The common theme of these areas is that they represent a radically new approach to problems that seriously limit propulsion technology. The technologies, however, differ in phenomenology and engineering approach. Active Fan and Compressor Stabilization One of the most troublesome phenomena in jet engine design and operation is compressor surge and stall. These are large-scale oscillations in air flow that result in abrupt thrust loss and can inflict severe mechanical damage. Current practice is to detune the compressor by 20–25 percent (with concomitant loss of performance) to avoid operating in regions where instabilities occur. Surge and stall also place significant restrictions on inlet design because distorted air flow from the inlets can trigger fan and compressor instabilities. FIGURE 8-4 Actively stabilized compressor suppresses rotating stall. In the past two years, laboratory-scale experiments have shown that it is physically possible to actively damp a compressor by using feedback control. The concept is to sense the disturbances when they are small and launch counter disturbances to damp them. The power required to do this is typically 10-5-10-6 of the compressor power. One such implementation, for the control of rotating stall in an axial compressor, is illustrated in Figure 8-4. Here, a circumferential array of transducers is used to detect small-amplitude traveling waves, which are damped by wiggling the compressor inlet guide vanes to launch counter waves. Decreases in the stalling mass flow of up to 25 percent have been demonstrated in this manner (see Figure 8-5).

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Aeronautical Technologies for the Twenty-First Century FIGURE 8-5 Active compressor stabilization moves the stall point to lower mass flows. Similar gains have been shown for the control of surge. There are significant potential advantages to stabilizing the fan and compressor against rotating stall and surge, including (1) lighter fan compressors (fewer stages and shorter chord airfoils), (2) improved compressor efficiency (more freedom in component matching), (3) increased distortion tolerance (shorter inlets, reaction control system considerations), (4) improved inlet-engine matching (reduced spillage drag), and (5) increased compressor design freedom. If the demonstrated gains are realizable in full-scale engines, system study has shown there can be dramatic improvement in aircraft performance or weight. Active Combustor Control Combustors also suffer from instabilities, especially in afterburners and ramjets. Active control of these instabilities has yielded both a 15-dB reduction in rumble instabilities in afterburner geometries and more complete combustion, permitting high power levels and shorter combustors. Active control has also been demonstrated in can-type geometries of main gas turbine combustors. Here, low-amplitude acoustic waves have been used to decrease the pattern factor

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Aeronautical Technologies for the Twenty-First Century (homogenize the flow) by 50 percent, potentially improving turbine durability and life. Active control may be necessary for lean-burn, ultralow-emission main burners for advanced subsonic and HSCT aircraft. Magnetic Bearings Magnetic bearings suspend the rotating members in magnetic fields, eliminating friction and lubrication requirements. Specific advantages over rolling contact bearings include elimination of the lubrication system, active damping of shaft dynamics and vibration, greatly increased temperature capability (up to 800°C), and large increases of load capability (two to four times). Magnetic bearings using conventional electromagnets are currently employed in large ground-based turbomachinery. Improved designs are capable of supporting typical gas turbine loads using less than 100 W. Since bearings of this type are inherently unstable, active control is required to maintain position. Active control enables the dynamics of the system to be optimized in software, greatly increasing design freedom and vibration damping. Current estimates are that an engine with magnetic bearings could have a 5 percent advantage in weight and a 5 percent advantage in efficiency over a conventional design. Active Noise Control Feedback control can be used to reduce noise either by direct wave cancellation or by influencing the noise-producing phenomena. Exhaust noise reductions of more than 20 dB have been demonstrated on large-scale ground-based gas turbines. Although perhaps further from practical aircraft realization than other smart engine technologies, active noise control offers the promise of a new approach to one of the most vexing of problems facing aviation, especially the HSCT. Recommendations for Smart Engines Should the results of these small-scale experiments extrapolate to full-scale engines, active control has the potential to improve aircraft propulsion performance and design in important ways. There is a long way, however, between small experiments and full-scale engine development. These are high-risk, high-reward technologies, with application unlikely before the turn of the century. As such, most of these concepts are beyond the horizon that U.S. industry is prepared to pursue. The Committee believes that NASA is uniquely positioned to pursue active control technologies for aircraft engines. NASA has both the personnel and the facilities to bring the technology along from university benches to the large-scale experiments necessary to assess concept viability prior to development by industry. NASA also has the breadth of disciplines at its centers and among its contractors to forge the teams (e.g., fluid mechanics, controls, structures) necessary to successfully pursue smart engine technology.

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Aeronautical Technologies for the Twenty-First Century TURBOMACHINERY COMPONENT TECHNOLOGY Significant advancements in the design methodology of turbomachines are needed to satisfy the requirements of both the advanced subsonic transport aircraft and the HSCT. Turbomachines for these systems should have fewer airfoils, reduced gaps between airfoil rows, lower aspect ratios, and higher clearance-to-span ratios. These machines must operate at high Mach numbers and lower Reynolds numbers, the latter in low-pressure turbines. The cooling air and turbine inlet temperatures are expected to be higher, whereas the amount of cooling and secondary air will decrease. The necessary solutions for these problems will emerge from the challenging fundamental research and development activity of the next 10–15 years. Continued advancements in CFD, together with improved flow measurement techniques, show the promise of utilizing more reliable analytical methods. However, the availability of advanced computational and experimental methods for turbomachinery design does not, in itself, ensure a more efficient and durable machine. The design process also relies very heavily on experience-based correlations and turbulence models and on design criteria. Implementation of advanced CFD codes without improvements in design criteria may not yield an improved design. The complexity of turbomachinery flows requires a national program to develop design systems for the next generation of gas turbine engines. Fluid Mechanics of Turbomachine Elements Among the issues facing turbine and compressor designers are the following: Two-dimensional airfoil optimization: Significant reductions in airfoil losses have been achieved over the past 15 years through the control of boundary layers, specifically by designing laminar flow and controlled diffusion airfoils. In current turbomachines, losses generated on airfoil surfaces constitute between 30 and 60 percent of total losses. Substantial increases in losses are measured for airfoils operating at transonic Mach numbers. Currently, low-loss airfoils in the transonic flow regime are designed by using simple criteria along with extensive experimental data. In the future, accurate numerical calculations from two-dimensional Reynolds-averaged Navier-Stokes codes can be used together with controlled experiments in plane cascades to develop design criteria for transonic flow applications. The guidelines needed to design low-loss airfoils for low Reynolds number application can be developed by using direct numerical simulation to identify transition criteria in separation bubbles. Endwall losses: The endwall regions contain 20–30 percent of the total losses in turbomachines. The physical mechanisms governing these losses are not well understood, and their magnitudes are 1.5–7 times larger than those estimated on the basis of wetted-surface calculations. There is a need to conduct accurate numerical simulations to identify loss generation mechanisms in the endwall region and to identify design criteria to develop low endwall loss passages.

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Aeronautical Technologies for the Twenty-First Century Tip clearance loss control: Tip leakage flows are responsible for 10–40 percent of the total losses in turbomachines. Although flows in the tip clearance regions are highly complex, the loss in efficiency due to tip clearance is linearly related to the fractional clearance. There is a potential to control tip clearance flows by designing airfoils to limit the amount of flow in the tip regions. Numerical simulations can play a very important role in this area if they are verified through experimental programs. Heat loads on airfoil pressure sides: Measured heat loads on the pressure sides of rotor airfoils are two to three times larger than those calculated by using current prediction methods. Since the amount of cooling air for rotor airfoils is controlled primarily by pressure-side heat loads, there is a need to develop reliable prediction methods to estimate them. Low heat load design concepts can potentially reduce cooling air requirements. Shock boundary layer/vortex interaction: Significant losses are generated from the interaction of shocks with boundary layers, and with tip leakage vortices. Shocks in fans cause the separation of boundary layers, which provides a conduit for the low-momentum flow to migrate toward the tip region. Shocks also interact with tip leakage vortices; this interaction can initiate stall in the rotor. Accurate numerical simulations must be conducted to identify the detrimental effect of shock interaction. There is also a need to develop turbulence models to provide accurate estimates of shock-induced losses on airfoil surfaces. Both numerical and experimental programs are required to develop flow prediction methods for high Mach number applications. Problems of Multistage Turbomachines Turbomachine flows are highly unsteady due to the relative motion of adjacent airfoil rows and incoming total pressure and total temperature profiles. Steady flow simulation methods, which have historically been used, fail to account for three specific flow features: (1) preferential migration of wakes from upstream airfoil rows; (2) preferential migration of hot and cold steaks toward the pressure and suction sides, respectively, of turbine rotors; and (3) preferential migration of endwall secondary flow and tip leakage vortices from the upstream airfoil row to the downstream airfoil passages, causing substantial effects on heat loads and losses. The effects of these flow phenomena are currently accounted for by using empirical correlations that are not very well grounded in flow physics. This situation will be further aggravated by current design trends. Available empirical correlations will not provide realistic flow behavior in the multistage turbomachines. Work needs to be initiated to develop physically sound models that account for the effects of periodic unsteadiness in multistage turbomachines. A promising multistage flow simulation strategy has been developed at LeRC, and concerted effort should be directed toward further development of this approach.

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Aeronautical Technologies for the Twenty-First Century Combined Computational Fluid Dynamics for Airfoil Rows Flow leakage from endwalls is expected to increase in future turbomachines. CFD codes, therefore, need to be developed and validated in order to provide accurate flow simulations for airfoil flow passages that have significant amounts of secondary air leakage. Prediction methods also need to be developed for airfoil rows having a significant amount of cooling air injection from airfoil and endwall surfaces. A rational approach is to develop a procedure that will compute flow through the internal passages of the turbine, a heat conduction code, and a solver for the airfoil passage. Such an approach is likely to provide a more accurate and faster method for predicting both performance and surface temperature for airfoil rows. Observations on NASA Turbomachinery Program The turbomachinery technology program at LeRC is extensive and contains some very significant and important elements, but, due largely to lack of funds and personnel, the pace of the work may be too slow to maintain the excitement it merits. It appears that the amount of innovative component research there is substantially less than it was two decades ago, when NASA contributed heavily in ideas and experimental results that led to strong advances in the industry. There are, however, excellent areas of work, of which three are mentioned below: Supersonic through-flow fan: This is a fine example of innovative technology firmly grounded in fundamental fluid mechanics, with the promise of significant performance gains in engines for supersonic flight. The design of the fan and the structuring of the experimental program have made effective use of the LeRC CFD capability, and its successful operation lends credibility to both the concept and the design method. Again, the pace of the work is slower than desirable or appropriate for a project carrying such promise. Essential work remains to be done in the design of the stator, improvement of the stage performance, and the mechanism of starting. Cooperative compressor research with Allison: LeRC has undertaken a cooperative research program with Allison on an advanced two-stage compressor with a pressure ratio better than 5:1, handling a through-flow of about 10 pounds per second. This is a high-performance compressor that is a very aggressive design and situates LeRC at the leading edge of compressor technology. Multistage turbomachine flow computation and design: This effort at LeRC is a highly innovative computational scheme realistic treatment of the fundamentally unsteady flow in multistage turbomachines. It represents an excellent example of scientific engineering, having its foundation in fundamentals while aiming at meaningful approximate calculations for design purposes. The technique recognizes the presence of fluctuations in the frequency range of blade passing and separately in the range of rotative speed, and works with relatively independent time averages in these ranges. To fulfill its promise, method requires experiments and computations focused on the determination of some of these average quantities; these are evidently not being pursued at present.

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Aeronautical Technologies for the Twenty-First Century These very promising activities suggest that increasing their number and augmenting their support merit serious consideration. Recommendations for Turbomachinery It has been the historic role of NASA, and one that industry has found beneficial, to provide a focus for research activity, to undertake innovative component experimental research, and to give leadership to teams consisting of industry, university and government agencies. The current environment requires, in addition, that NASA assist in providing computer resources for numerical simulation and that it cooperate in the development and standardization of various CFD codes to facilitate evaluation against basic data and permit ready incorporation into design codes. NASA has been slow to respond to the exciting potential of stall and surge control and has not yet taken steps toward development of sensors and actuators required for its implementation. Although the capability of such control has been unequivocally demonstrated, the development of techniques appropriate to operational engines demands a strong and innovative effort. Another area in which additional effort would be welcome is the problem of short blades that will be encountered in latter stages of new, very high pressure ratio core engines. NASA's work on the off-axis independent compressor is very promising but should not constitute its total effort in this area. COMBUSTORS AND EMISSIONS Although the emission of nitrogen oxides, carbon monoxide, and unburned hydrocarbons constitutes a significant issue for subsonic commercial transports, and promises to become a larger problem for short-haul aircraft as this traffic increases, the currently intense interest centers on the HSCT. Here, the NOx emissions are a central issue for flight in the stratosphere, where particulate exhaust and the formation of ice crystals together with the oxides of nitrogen constitute a potential threat to the ozone layer. The development and design issues, which are already sufficiently complex, are further clouded by currently inadequate, but slowly evolving, models of the upper atmosphere that will play a role in setting standards for altitude exhaust emissions. Moreover, it is not clear at present just how a known body of engine exhaust information would be coupled meaningfully to the grid of atmospheric models now in use. Engines for current subsonic transport aircraft emit between 10 and 20 g of nitrogen oxides per kilogram of fuel (see Figure 8–6), depending on whether the combustors are of the conventional design or the more recent dual annular design. Under takeoff conditions, emissions from these same combustors increase to between 20 and 30 g/kg fuel. If this performance is scaled on the basis of the nitrogen oxide ''severity index'' currently employed, the HSCT would be expected to produce between 30 and 40 g/kg fuel at cruise conditions. The goal for HSCT

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Aeronautical Technologies for the Twenty-First Century FIGURE 8-6 Nitric oxide emissions versus severity index(s), current and future engines. engines is 5 g/kg fuel—between 12 and 16 percent that of current engines—the requirement could be set even lower, more out of uncertainty than rationality. Clearly this reduction by a factor of six or more requires heroic measures. Recognizing these facts in the early 1970s, LeRC embarked on a low-emissions burner program that focused on what became known as the lean premixed prevaporized burner. Subsequently, this effort led to NASA sponsorship of separate research and development programs with P&W and GE. Two candidate burner technologies were identified and these continue to be pursued by the two engine manufacturers. Research on the chemistry and gas dynamics of NOx reduction continues at LeRC. The production of nitrogen oxides increases with pressure, temperature, and residence time in the combustor. Consequently, reduction techniques focus on just what latitude one has with these variables in view of the engine cycle requirements. Not only is temperature the most sensitive of these, but owing to the design of conventional burners, it also offers the greatest possibility for control. In the conventional engine, fuel is initially burned at approximately stoichiometric conditions and subsequently diluted to the desired leaner condition. The high temperatures in the primary combustion zone result in rapid production of NOx during its residence time and set the value of the final emission level. The advantages of this arrangement are that the hot, stoichiometric primary zone provides good stability, ignition, and relight, while the addition of dilution air allows convenient cooling of the combustor liner. The low-NOx burners are consequently designed to avoid the hot stoichiometric and dilution zones, thereby reducing emissions, but at the expense of stability and cooling problems.

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Aeronautical Technologies for the Twenty-First Century The NOx formation rate shows a strong dependence on mixture ratio, as a result of the dependence of temperature on mixture ratio. In a conventional burner the primary region produces NOx at nearly the peak value, and subsequent air dilution decreases the production rate to less than 1 percent peak value at combustion discharge conditions. The lean premixed, prevaporized concept reverses the procedure by preparing the fuel/air mixture at its desired final value and carrying out the combustion at a very low NOx production rate. The near symmetry of the NOx production rate about the stoichiometric mixture suggests the alternative procedure of burning with a rich mixture and subsequently diluting it rapidly to the desired final lean condition. This technique underlies the second type of low-NOx burner under serious consideration. A combustor that introduces a lean, premixed, prevaporized mixture into the combustion chamber is being pursued by GE. With this design it is necessary to mix the air and the fuel vapor thoroughly on the molecular scale to achieve the indicated reduction of nitrogen oxides. Moreover, because this mixture volume constitutes a source of preignition and flame flashback problems, the mixing must be very rapid to minimize the mixture volume and residence time. To accommodate the anticipated range of operating conditions, variable-geometry vapor inlet and bypass air controls are required as well as the corresponding sensors. Furthermore, because the larger part of the air is mixed before combustion is carried out, the dilution air conventionally available for combustor liner cooling is much less, accessible. As a consequence, high-temperature materials are essential. The second option, that of rich burning, is being pursued by P&W and, for reasons that will become obvious, has been designated the Rich Burn Quick Quench burner. In this burner the fuel and a limited amount of air are introduced into the rich stage. Here again, it is absolutely essential that the fuel-air mixing be carried out on a molecular scale before combustion occurs. To the extent that mixing is nonuniform, the reaction will preferentially take place in those regions where the mixture is close to stoichiometric, thus partially defeating the aim of the burner. As dilution air is added, each portion of the mixture passes through the stoichiometric ratio and will produce nitrogen oxides at that very rapid rate. Consequently, rapid mixing during the quench phase is essential. Again, because the temperatures are higher than conventional and because bleed liner cooling is unacceptable in the rich burn stage, high-temperature materials are essential. These ultralow-NOx burners entail higher than conventional pressure losses due to the requirement for rapid mixing, the use of new high-temperature materials, variable geometry, and the attendant controls and sensors. In addition, it appears that these burners may be somewhat longer than conventional and entail an increased engine weight that is, as yet, undetermined. The cooperative development of the HSCT engine, between P&W and GE, permits efficient and economical exploration of these two burner types; it is essential that NASA, P&W, and GE get their Joint Technology Acquisition Program in place. In addition, it is of the utmost importance that LeRC continue vigorously its basic research in this area, driving toward NOx levels even lower than the present goal of 5g/kg fuel. Airport NOx emissions will be a challenge for the advanced subsonic engine cycles with 70:1 pressure ratio and 3000ºF turbine inlet temperature. With current combustor types, the

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Aeronautical Technologies for the Twenty-First Century NOx per unit of thrust could be tripled under these conditions. A good target would be to decrease NOx emission to an absolute level half that of current engines. This will require very advanced combustors employing variable geometry, new materials, and smart controls.

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