Distributed Turboelectric Propulsion for Hybrid Wing Body Aircraft

Distributed Turboelectric Propulsion for Hybrid Wing Body Aircraft

Hyun Dae Kim, Gerald V. Brown, and James L. Felder NASA Glenn Research Center Cleveland, Ohio, USA

Abstract

Meeting future goals for aircraft and air traffic system performance will require new airframes with more highly integrated propulsion. Previous studies have evaluated hybrid wing body (HWB) configurations with various numbers of engines and with increasing degrees of propulsion-airframe integration. A recently published configuration with 12 small engines partially embedded in a HWB aircraft, reviewed herein, serves as the airframe baseline for the new concept aircraft that is the subject of this paper. To achieve high cruise efficiency, a high lift-to-drag ratio HWB was adopted as the baseline airframe along with boundary layer ingestion inlets and distributed thrust nozzles to fill in the wakes generated by the vehicle. The distributed powered-lift propulsion concept for the baseline vehicle used a simple, high-lift-capable internally blown flap or jet flap system with a number of small high bypass ratio turbofan engines in the airframe. In that concept, the engine flow path from the inlet to the nozzle is direct and does not involve complicated internal ducts through the airframe to redistribute the engine flow. In addition, partially embedded engines, distributed along the upper surface of the HWB airframe, provide noise reduction through airframe shielding and promote jet flow mixing with the ambient airflow. To improve performance and to reduce noise and environmental impact even further, a drastic change in the propulsion system is proposed in this paper. The new concept adopts the previous baseline cruise-efficient short take-off and landing (CESTOL) airframe but employs a number of superconducting motors to drive the distributed fans rather than using many small conventional engines. The power to drive these electric fans is generated by two remotely located gas-turbine-driven superconducting generators. This arrangement allows many small partially embedded fans while retaining the superior efficiency of large core engines, which are physically separated but connected through electric power lines to the fans. This paper presents a brief description of the earlier CESTOL vehicle concept and the newly proposed electrically driven fan concept vehicle, using the previous CESTOL vehicle as a baseline.

Nomenclature I. Introduction

AC alternating current According to a number of air traffic forecast studies, the BLI boundary layer ingestion growth in air travel in the United States or world will increase BWB blended-wing-body by a factor of 2 to 4 by 2025 (refs. 1 and 2). This continued CAEP Committee on Aviation Environmental growth in the passenger and freight air traffic will require Protection better utilization of available airport assets. Large airports with CESTOL cruise-efficient short take-off and landing long runways (>10 000 ft, 3050 m) are already heavily utilized EBPR effective bypass ratio (ratio of mass flow rate while small airports with runways too short (<3000 ft, 910 m) through all fans to rate through engine core) to support large transport class jet aircraft are often EIS entry into service underutilized. Table 1 shows a number of metropolitan FAR Federal Aviation Regulations airports around 15 major U.S. metropolitan areas with at least HTS high temperature superconducting an intermediate size runway length of 3000 ft (~910 m). Most HWB hybrid wing body of these cities have at most one or two large airports handling hp horsepower (1 hp ~ 0.7456 kW) much of their large transport aircraft traffic, but they also have IBF internally blown flap additional regional airports nearby with shorter runways to IOC initial operating capability accommodate smaller aircraft. For example, the city of Atlanta LTO landing and take-off has one large-capacity, long-runway airport within the city PAI propulsion airframe integration boundary but has four more regional airports with at least SFW subsonic fixed wing 3000 ft runway lengths within 20 miles (~32 km) of the city STOL short take-off and landing metropolitan area.

TSFC thrust specific fuel consumption TOGW take-off gross weight USB upper surface blowing TABLE 1.—NUMBER OF U.S. METROPOLITAN (METRO) AIRPORTS TABLE 2.—NASA’S TECHNOLOGY GOALS FOR FUTURE SUBSONIC WITH AT LEAST 3000-ft (~915-m) RUNWAY LENGTH AROUND 15 FIXED WING (SFW) VEHICLES a a a MAJOR METRO AREAS Corners of the N +1 (2015 EIS) N +2 (2020 N +3 (2030– Metropolitan areas Number of metro airports within 20 trade space Generation IOC) 2035 EIS) miles ( ∼ 32.19 km) Conventional Generation Advanced Tube and Wing Conventional Aircraft Atlanta 5 (relative to Hybrid Wing Concepts Charlotte 5 B737/CFM56) Body Chicago 5 (relative to Houston 9 B777/GE90) Las Vegas 4 Noise –32 dB –42 dB 55 LDN at Los Angeles 11 (cumulative average Minneapolis 6 below airport New York 7 Stage 4) boundary Philadelphia 8 LTO NOx –60% –75% Better than Phoenix 8 Emissions –75% San Francisco 4 (below CAEP/6) San Diego 4 b b Performance: –33% –40% Better than Seattle 7 Aircraft Fuel –70% South Florida Burn Miami 5 Performance: –33% –50% Exploit c Orlando 4 Field Length metroplex Tampa 8 concepts a Washington-Baltimore 8 “N” represents current state-of-the-art aircraft as stated in parentheses.

b Number of U.S. airports in 15 108 An additional reduction of 10 percent may be possible through improved metro areas operational capability.

c Concepts that enable optimal use of the airports (with shorter runways) within the metropolitan areas.

In order to meet future traffic demand with limited airport access, revolutionary airplane concepts are needed that can from regional airports (refs. 3 and 4). The configuration for utilize these smaller airports. For these new concepts to be that study utilized 12 small conventional high-bypass-ratio successful, they must dramatically reduce take-off and landing turbofan engines, each with about 7000 lb (~31 000 N) of noise, due to the urban setting of many of these fields, and yet thrust at sea level, powering a HWB vehicle. The HWB is the still carry an economically viable number of passengers and main object of study to meet NASA’s N+2 goals. Because the freight over transcontinental distances at current jet transport results of that study are newly published, and have not been speeds. At the same time, these new aircraft must dramatically widely disseminated, they are summarized in the next section reduce energy consumption and environmental impacts. In for background and provide the baseline for the current study.

response to growing aviation demands and concerns about the Recently, a very low noise “Silent Aircraft,” based on the environment, NASA’s Subsonic Fixed Wing (SFW) project blended-wing body or BWB airframe and on distributed identified four “corners” of the technical trade space—noise, propulsion, was proposed and studied. Its objective was to emissions, aircraft fuel burn, and field length—for aircraft contain objectionable noise within the airport boundary and to design. Table 2 lists these technology goals for three future improve vehicle fuel efficiency (ref. 5). This configuration had timeframes, where N+1, N+2, and N+3 represent the years a number of new technologies, including embedded turbofan 2015, 2020, and 2030, respectively. Although it may not be engines with each engine core driving three fans through a feasible to meet all the goals for each timeframe, the multi- gear and shaft system, yielding a very high bypass ratio objective studies will attempt to identify possible vehicle (ref. 6). The increased engine bypass ratio provided both low concepts that have the best potential to meet the combined thrust specific fuel consumption (TSFC) and low engine noise.

goals.

To improve vehicle performance enough to meet NASA’s One of the vehicle and propulsion concepts that NASA is N+3 goals, a drastic change in propulsion system is required.

exploring for N+2 is a synergistic combination of a hybrid A newly proposed vehicle, which is the subject of the present wing body (HWB) airframe and a distributed propulsion paper, uses the baseline cruise-efficient short take-off and system. A number of fixed wing aircraft using “distributed landing (CESTOL) aircraft airframe mentioned above but propulsion” have been proposed and flown before, although employs superconducting motors to drive the distributed fans what constitutes distributed propulsion is not clearly defined.

rather than a number of small conventional high bypass ratio Examples include the 1940’s YB–49 flying wing aircraft with engines. The power needed for these electric fans comes from four completely embedded engines in each side of the wing two remotely located gas-turbine-driven superconducting and the 1960’s Hunting H.126 jet flap research aircraft, which generators through electric power lines. This arrangement diverted almost 60 percent of its thrust across its wing trailing allows many small partially embedded fans while retaining the edge to achieve very high lift capability.

superior efficiency of large core engines. The next section NASA funded a 1-year study that evaluated the synergistic presents a brief description of the baseline CESTOL vehicle benefits of distributed propulsion and airframe integration and propulsion concept followed by the newly proposed with respect to cruise efficiency and quiet operation of aircraft electrically driven fan concept vehicle.

To address the CESTOL vehicle with low noise

II. Distributed Propulsion Concepts

characteristics, a CESTOL vehicle configuration was developed jointly by NASA and Boeing to utilize short A number of distributed propulsion vehicle concepts for the runways at regional airports and was reported by one of the HWB platform have been studied recently (refs. 3 to 8). The authors (ref. 3). A brief summary of this concept is described motivation has been to increase aircraft performance, to lower in the first subsection below. Combining that earlier CESTOL the noise to the surrounding community, and/or to enable short vehicle concept with our presently proposed turboelectric take-off and landing (STOL) capability. The following propulsion system results in the current conceptual study, possible benefits of distributed propulsion HWB configuration which is designated “Turboelectric-Powered CESTOL compared to a conventional “tube-and-wing” configuration Concept” and is described in the second subsection below.

have been identified: • Reduction in fuel consumption by ingesting the thick Baseline 12-Engine CESTOL Concept boundary layer flow and by filling in the wake generated by the airframe with the engine thrust stream (refs. 9 to 11).

To develop a distributed propulsion CESTOL vehicle • High lift via high-aspect-ratio trailing-edge nozzles for configuration, NASA and Boeing performed a joint study, vectored thrust providing powered lift, boundary layer which we summarize in this section. The initial configuration control, and/or supercirculation around the wing, all of was based on the HWB because of its high cruise efficiency, which enable short take-off capability (refs. 12 and 13).

low noise characteristics, and a large internal volume for • Reduction in aircraft noise to the surrounding community integrating embedded distributed propulsion. The preliminary through airframe shielding (refs. 3 and 4).

vehicle analysis is reported by Kawai (ref. 16) and • Improvement in safety through a redundant propulsion summarized in reference 3, and the vehicle configuration is system.

shown in figure 1. The powered lift system was selected • Reduction in aircraft propulsion installation weight through because of the high lift efficiency of the internally-blown-flap inlet/nozzle/wing structure integration.

(IBF) concept. A distributed propulsion system with 12 • Elimination of aircraft control surfaces through differential conventional turbofan engines with approximately 7000 lb and vectoring thrust for pitch, roll, and yaw moments.

(~31 000 N) thrust each would enable high lift by using low • High production rates and easy engine replacement of pressure fan bypass air that would not have hot duct issues and engines that are small and light.

that would be subsonic to keep the powered lift noise down.

The concept here is the use of distributed embedded • Application of nontraditional engine concepts such as the multifan engine or electric fans. propulsion for quiet IBF powered lift with substantial engine noise shielding, including some jet noise shielding. The The large available volume in the HWB configuration may CESTOL concept combines substantial engine noise shielding facilitate use of hydrogen or other alternative fuels to achieve zero or near zero emissions (refs. 14 and 15).

with rapid climb out and steep descent to provide a very low Turboelectric-Powered CESTOL Concept noise footprint. The preliminary noise analysis of the vehicle To meet the aggressive NASA SFW N+3 goals in table 2, is reported by Stone (refs. 4 and 17).

we have begun a study that carries over the baseline CESTOL Based on current trends in air transports and STOL airframe, but we propose a more radical propulsion system considerations, the following mission requirements were used that replaces the discrete turbofan engines. We propose a for the vehicle: turboelectric propulsion system with superconducting electric • Payload: 40 000 lb (~18 000 kg) fans powered by two turbine-engine-driven electric generators.

• Range: 3000 nm (~5.600 km) A notional vehicle is shown in figure 2. Because this new • Speed: Mach 0.8 at 30 000 ft (~9000 m) effort focuses on the propulsion system, the airframe has not • Field length: <5000 ft (<1500 m, FAR Part 25) been reexamined in light of the new propulsion system.

Therefore, an airframe similar to that of the earlier CESTOL • Climb at Std + 15 ° C configuration was retained as a baseline and the distributed • Landing flare for passenger comfort with a 6 ° glide slope electric propulsion system was applied instead of discrete With these requirements and using Boeing’s WingMOD small turbofan engines.

(ref. 18) multidisciplinary optimization code, an aero- The initial propulsion system consists of two wing-tip- dynamically trimmed vehicle configuration was obtained and mounted turboelectric generators and a set of 16 small electric mission performance data were determined. The following is fans. The 35-in.- (~90-cm-) diameter fans are distributed along the set of predicted vehicle and performance parameters: a large portion of the upper aft wingspan to maximize the benefits of boundary layer ingestion (BLI). The number of • Take-off gross weight: 189 140 lb (~85 792 kg) fans was chosen on the basis of assumed available span width, • Total fuel: 44 098 lb (~20 000 kg) nacelle length, and inlet and nozzle geometry constraints. To • Take-off field length: 2452 ft (~747 m) increase BLI benefits and to minimize interference drag between the fan and external flows, contiguous “mail-slot” • Take-off C = 1.66 Lmax inlets, high-aspect-ratio slot nozzles, and span-wise- • Initial cruise altitude: 39 000 ft (~11 887 m) continuous upper nacelles were adopted. Five outboard low- • Landing field length: 3477 ft (~1060 m) pressure-ratio fans on each side of the vehicle are used for • Landing C = 1.06 Lmax powered lift and six center fans are used as pitch effectors at The take-off field length is for obstacle clearance with an take-off rotation. For producing powered lift, upper surface blowing (USB) is deemed to be better than internally blown engine out. However, because many engines (12) were distributed on the wingspan, the engine-out condition did not flap (IBF), because of structural and mechanical simplicity.

Based on the baseline 12-engine CESTOL concept thrust include lateral control drag because only one engine inoperative (out of 12) would produce no significant yawing requirement, the total shaft power for the vehicle at sea-level static conditions is assumed to be approximately 84 000 hp moment at a mission-critical stage (mainly at takeoff). Indeed, aircraft with powered-lift distributed-propulsion systems may (horsepower (63 MW)) and the total available shaft power at cruise is assumed to be 25 000 hp (19 MW), which require a general reexamination of engine-out airworthiness certification regulations because controllability limits are corresponds to approximately 1500 hp (~1.1 MW) for each fan at cruise. The wing-tip-mounted engine-core/turboelectric currently based on one engine out. Note that the landing field length is about 3477 ft (~1060 m), which includes the 1.67 generator is also analyzed and the estimated effective bypass ratio (EBPR, ratio of mass flow rate through all fans to rate factor on stopping distance. It is believed that the use of a variable area nozzle for improved powered lift during through engine core) for the whole propulsion system is approximately 10, which is higher than that of present approach would enable further reductions in field length.

Embedded distributed propulsion enables the use of low- turbofan engines (and of the 12-turbofan system), promoting fuel efficiency.

pressure fan-bypass air for an IBF system, wherein a high- aspect-ratio slot nozzle is used in conjunction with a slotted Although this kind of distributed propulsion concept, with a small number of turboelectric generators driving numerous airfoil with the nozzle exhaust pumping through the slot to increase circulation and lift. The small diameter engines with a electric fans, could be applied to other vehicle architectures (e.g., conventional tube and wing aircraft), the concept is bypass ratio of 9.4 have forward noise shielding and employ mixer nozzles to increase the jet noise frequency and move the perhaps most naturally applied to the current CESTOL vehicle configuration to reduce fuel consumption, noise, emissions jet noise source locations forward. The forward jet source noise can then be shielded by airframe surfaces to reduce aft and field length as noted before. Nevertheless, the following are identified as possible advantages of using a turboelectric and sideline noise. A more complete description of noise analysis methods and results can be found in reference 17. drive system on an arbitrary “platform”: at a reduced but symmetric thrust with the electric power • Decoupling of the propulsive device from the power- from the remaining turbogenerator using a common bus producing device. This is the major departure from the network.

current state-of-art aircraft vehicle/engine design, possibly enabling unprecedented performance and design flexibility • Asymmetric fan thrust available for yaw control because of of the air vehicles. The turbine-engine-driven generators and the fast response electric motors. The fan power, and hence the electric-motor-driven fans can be located at their thrust, on one side of the aircraft can be increased and that optimum locations in the aircraft to maximize total vehicle on the opposite side reduced, keeping the total power from performance and operation. the turboelectric generators constant. Thus, the total thrust can remain constant while yawing the vehicle.

• High fuel efficiency due to high EBPR. EBPR is defined as the ratio of mass flow through all fans to the mass flow • Use of alternative fuel, for example, hydrogen or electrical through the engine cores. power sources such as fuel cells. Cryogenic hydrogen, used as fuel, could provide the required cooling to maintain • Speed of the power turbine shaft in the turbine engine independent of the propulsor shaft speed—the electrical superconductivity in the electric generators and motors.

system functions as a gearbox with an arbitrary gear ratio. • Large electrical power off-take capability for in-flight and With the addition of power electronics, the two shaft speeds ground use.

can change independently, giving the effect of a variable ratio gearbox. This allows the shaft speed of the power Furthermore, the following are identified as possible turbine in the core engine to be optimized without the usual “vehicle specific” advantages for the currently proposed concern that a low fan pressure ratio requires a low engine propulsion concept: shaft speed (because of blade tip speed constraints), which increases the size and weight of the power turbine in direct- • Lower TSFC with large engines and electrically driven drive turbofan engines.

multiple fans than with the multiple small turbofans used in • Minimal engine core jet noise due to maximum energy the baseline CESTOL configuration extraction to provide power to remotely located fans. • Higher propulsive efficiency via continuous spanwide • Symmetric thrust in the event of a turbine engine or boundary layer flow ingestion and wake fill-in with the fan generator failure. All fan modules could continue operating thrust stream flexibility to the propulsion and vehicle architecture. As noted • Direct powered lift through continuous spanwide USB using low-pressure fan air above, the electric components function as a gearbox allowing the turbine engines to run at high speed, independent of the • Very low community noise using low-pressure ratio fans fan-shaft speeds. Beyond functioning as a simple gearbox, the and airframe shielding electric components can function as a continuously-variable- • Minimal engine rotor blade burst impact on passengers and ratio gearbox with the addition of a solid-state converter. This vehicle structure due to the wing-tip location of the would permit the turbine engines to run at the most effective turboelectric generators and the numerous small fans shaft speed regardless of the required changes in the fan-shaft mounted on the rear top side of the vehicle speed as airspeed, altitude and noise limits change. Higher • Reduction of lift-induced drag and of wake vortices due to part-load efficiency can thereby be achieved. However, in this the wing-tip location of the engine cores (ref. 19) initial paper, we consider only a fixed speed ratio, which can • High engine core inlet pressure recovery similar to be achieved without using a solid-state converter. Figure 3 conventional aircraft podded engine installation illustrates the components of a turboelectric propulsion • Lower propulsor nacelle structural weight due to absence of system, including the optional power converters.

sudden internal pressure rise (hammer shock) from engine core stall Superconducting Motors and Generators • Use of conventional low-temperature material on thrust vectoring mechanism due to “cold” fan air discharge Superconducting materials lose all their electrical resistance • Lower wing structure weight through better load below a “critical” temperature and can carry high current in distribution with wing-tip-mounted engine core and small wires or tapes, leading to light, compact, very efficient distributed span-wide fan installation (ref. 20) motors and generators. The operating temperature required for • Low cabin noise due to remote location of engines and superconducting windings is somewhere between 20 K (the propulsors away from the passenger cabin area normal boiling point of liquid hydrogen) and 65 K (somewhat • Easier maintenance access to the gas turbine and electric below the normal boiling point of liquid nitrogen). The state- generator than with the embedded engine configuration of-the-art of cryogenic and superconducting motors and generators is reviewed in references 22 to 24. Machines as However, using a distributed turboelectric propulsion large as 35 MW output (ref. 25) and as fast as 15 000 rpm system with superconducting devices may present adverse have been tested or designed. The higher performance effects in overall vehicle performance and operation. The machines are intended for military applications, but prototypes following are identified as possible drawbacks of the for commercial machines are beginning to appear. High- electrically driven system and of the newly proposed vehicle: temperature superconducting (HTS) machines for aircraft • Weight increase due to core generators, motors, and balance propulsion have previously been discussed, primarily with of the superconducting system “tube-and-wing” aircraft in mind (refs. 21 and 26).

• Possible nonlinear aircraft control laws due to interactions For turboelectric aircraft propulsion, motors and generators between the external aerodynamics and the propulsion with HTS windings on both the rotors and the stators are system envisioned. In most state-of-the-art machines that are called • System complexity due to additional new technology superconducting today, only the rotor windings are • Operational difficulties with superconducting parts and superconducting. The stator windings, which are the high- cryogenic fluids power windings where most of the losses occur, are made of • Ice, snow, rain, etc., ingestion by the upper surface BLI copper and operate at room temperature. As of this writing, only a few small experimental machines have been made with mail-slot inlet superconducting stators. The reason is that, whereas the rotor It will be necessary to use superconducting motors and carries direct current and dissipates little power, the stator generators rather than conventional motors and generators in carries alternating current (AC) and has losses that depend on the aircraft propulsion system to reduce the weight fraction of the fineness of the filaments in the superconducting composite the propulsion system. Conventional electric generators and wire in the winding. Reducing those losses requires some motors are far too heavy to be used on a large transport air technology development. It appears reasonable that the AC vehicle (ref. 21). A description of the superconducting system losses in a superconducting stator can be reduced to less than is presented in the next section.

0.1 percent of the machine’s output power, with a developmental goal as low as 0.01 percent.

The electric power would be carried from the generators to

III. Superconducting Electric Drive

the motors by HTS transmission lines. Such lines are presently

System

being tested in the electric grids of congested urban areas.

They can carry hundreds of megawatts of power with less than The use of gas-turbine-driven generators to supply electric 10 kg/m of mass and only a few W/m of loss (ref. 27).

power to motor-driven propulsive fans adds considerable achieved, it is by no means assured. Inadequate developments Refrigeration Options of lightweight cryogenic refrigerators would make tanked The low temperatures required for the electrical components liquid hydrogen the preferred cooling option. If can be viewed as analogous to the lubrication required for a superconductors with sufficient AC tolerance are not gearbox. Both are required to remove waste heat from the developed, then turboelectric propulsion could still be an power transferring components. In the electrical case, this option on liquid-hydrogen-fueled aircraft.

removal might be achieved in three different ways. If future aircraft are liquid hydrogen fueled, that fuel can cool the Preliminary Weight, Efficiency, and Performance electrical components before being burned in the turbine Estimates engines. That refrigeration mode entails minimal weight or efficiency penalty and could enable turboelectric propulsion In spite of uncertainty of the future level of refrigerator and even with motors and generators that employ cryogenic, but AC superconductor technology, we present some weight and nonsuperconducting, pure-metal conductors in their stators.

efficiency estimates that are based on the level of development On a purely jet-fueled aircraft, refrigerators must be used.

that we expect for all-superconducting generators and motors.

They may be major components, depending on the efficiencies Weights as a function of power, based on electromagnetic and and on the weight per input power of the refrigerator. As loss analyses from references 28 to 30 and structural weight discussed below, reasonable technology developments are estimates, are shown in figure 4. Optimization was performed required to make the refrigeration system manageable. An to minimize motor (or generator) weight plus refrigerator alternative for primarily jet-fueled aircraft is to carry only weight. The refrigerator, with our assumptions, weighs enough liquid hydrogen (with a reserve) to cool the electric ~70 percent as much as the motor or generator that it cools.

system and then to use the hydrogen as fuel, so it would Efficiencies, including the refrigerator power, are at least contribute to the aircraft’s total fuel complement. The stored 99.4 percent. Figure 4 shows that the expected weight of a liquid hydrogen would represent less than 10 percent of the motor or generator with its cooler is considerably less than the total fuel heating value on the aircraft, if the electrical losses weight of a turbine engine core for equal power. Weight and are reasonably low.

efficiency comparisons are made in table 3 among three The weights and efficiencies of the electrical components of propulsion systems: a 16-fan turboelectric propulsion system, a turboelectric propulsion system will depend strongly on the 16 independent small turbofan engines, and 2 large level of technology development over the next 20 years or so conventional turbofans. The core engine and generator in the (especially with respect to cryogenic refrigerators and AC turboelectric system were uprated 0.9 percent to 42 380 hp to tolerant superconductors). While there is reasonable basis to compensate for the 1 percent loss in the electric system at expect that the assumed technology development can be TABLE 3.—COMPARISON OF DIFFERENT PROPULSION SYSTEMS take-off conditions, but the core weight estimate was reduced Propulsion Components Weight, lb Efficiency, TSFC, (from the value typical for a turbofan engine) because of an –1 System (kg) % hr assumed one-third reduction in low-pressure shaft and power Two 42 380- 7300 ---- 0.57 Turboelectric turbine weight due to higher shaft speed and a shorter power hp engine (3300) distributed fans cores (refrigerated) shaft. The turboelectric system weighs 5000 lb (2300 kg) more Two 42 380- 3000 99.7 ---- than the 16-engine system but has 9 percent lower TSFC hp electric (1300) including the 1 percent electrical and refrigeration loss at generators takeoff. (The electrical loss at cruise would be lower, due to (including refrigerators) much lower required refrigerator power at cruise, but the take- Sixteen 4700 99.4 ---- off efficiency is used to make the estimates that follow.)

5250-hp (2100) Table 3 shows a comparison of different propulsion systems motors (16 turboelectric distributed fans with 2 cooling options, 16 (including conventional small turbofan engines, and 2 conventional large refrigerator) Total 15 000 99.1 ---- turbofan engines) with the same thrust requirement. Weights (6800) exclude propulsors (fans), which would have similar total Two 42 080- 7300 ---- 0.57 Turboelectric weights in all systems. TSFC values shown in table 3 are hp engine (3300) distributed fans based on best present-day values for the engine size.

cores (LH cooled) Refrigerator weight is based on 5 lb per hp input and Two 42 080- 1900 99.9+ ---- hp electric (860) 30 percent efficiency and HTS AC losses on a 12- μ m filament generators characteristic dimension.

(LH cooled) To compare the turboelectric system with the 16-turbofan Sixteen 3100 99.9+ ---- system, by balancing out the opposite effects of lower SFC and 5250- hp (1400) motors (LH higher weight of the turboelectric system, the Breguet range cooled) equation, sufficient to determine relative ranking, is applied to Total 12 300 99.9 ---- both systems, with the requirement of equal aircraft range and (5600) a approximating the entire flight as cruise. Solving for the Conventional Sixteen 10 000 91 0.63 small 5250-hp (4500) required change in fuel weight between the 16-engine case and distributed engine cores the turboelectric case, we find that the turboelectric aircraft turbofans would require 7 percent, or 3000 lb (1400 kg) less mission fuel.

Conventional Two 42 000- ---- 8700 0.57 Thus, the slightly heavier turboelectric aircraft would have a net large hp engine (4000) nondistributed cores fuel savings of roughly 7 percent on each flight, compared to turbofans the baseline aircraft powered by 16 small engines. This estimate a Relative to 42 000-hp engine core at 0.57 thrust specific fuel consumption.

will be refined as the study progresses with a detailed mission aspects of the superconducting motors and generators, analysis. Known omissions in the weight estimates of the development is required on subsystems and auxiliary systems.

electric system include the superconducting transmission lines The largest potential technology development payoff is in (estimated at only 3 percent of the turboelectric system reducing the AC losses in HTS motors and generators. Those weight) and other power management and distribution losses must be well below 1 percent in each machine to keep components. the required refrigeration reasonable. (Note that large If the motors and generators were cooled by liquid hydrogen generators already exceed 99 percent efficiency, even at room (with only enough carried on the aircraft to provide temperature.) The several types of AC losses that occur in refrigeration) rather than refrigerators, then the turboelectric HTS materials can be reduced by reducing the size of HTS system would weigh 2300 lb (1000 kg) more than the 16- filaments in the composite conductor and twisting them. An engine system, and the required jet fuel is reduced by 4000 lb order of magnitude or more decrease in size from present (1800 kg), or 9 percent (calculated from the efficiency practice is required. Such dimensions (and smaller) have been advantage of the large engines, without accounting for the achieved in the older low-temperature superconductors, replacement of jet fuel energy with liquid hydrogen energy), indicating promising approaches for the newer HTS materials.

and TOGW drops by 560 lb (255 kg). This estimate does not In addition, the required refrigeration is proportional to the include corrections for the weight of the liquid hydrogen above losses, as is the required input power to drive the (which would provide about 5 percent of the aircraft’s fuel refrigerator and hence the refrigerator weight. Present energy) and its tankage and accessories, compared to the cryogenic refrigerators of the required capacity have not been corresponding weight reduction of the jet fuel, tankage, and designed with low weight as an objective and must reach components. (It may be noted that, for the same energy, liquid significantly lower weight per input power to be acceptable on hydrogen has almost 4 times the volume but only one-third the aircraft. A factor of 3 to 6 reduction from the present best weight of jet fuel.) machines is desired. Improvements in refrigerator mechanical A comparison between the turboelectric case and two large efficiency would also be effective but may be more difficult to (presumably podded) turbine engines can be made based on achieve. As noted above, no refrigerators would be required the numbers in table 3. One can see that the entire refrigerated on liquid-hydrogen-fueled aircraft or on ones carrying enough turboelectric system weighs 6300 lb (2900 kg) more than two liquid hydrogen inventory to cool the electric components.

large turbofan engine cores of 42 000 hp each (with no weight A wide range of analyses and system studies would be allowance for podding) and would be ~1 percent less efficient beneficial. To determine the optimal fan pressure ratio and at takeoff because of the electrical losses. A liquid-hydrogen- other propulsion system parameters, a detailed mission cooled turboelectric system would weigh 3600 lb (1600 kg) analysis is needed, which would include optimizing both the more than the large turbofan engine cores. Thus, the fan propulsor modules and the thermodynamic cycle of the propulsion system weight for an HWB using podded engines engine. Other propulsor options, such as ducted-propeller would be significantly less than either of the two turboelectric systems, should be examined. In addition, the basic mission systems discussed, with consequent accompanying reductions profile needs to be examined to determine the impact of cruise in fuel burn. However, the use of two separate podded engines Mach number on mission fuel burn, block times, and direct would provide no STOL capability and only limited noise operating costs for different fuel prices. The unique flexibility reduction, two important corners of the trade space, and none of the turboelectric propulsion system is well suited to the of the other potential benefits and capabilities mentioned examination of a wide range of propulsion and mission above. options.

IV. Further Study and Research V. Concluding Remarks

Directions

Two novel transport vehicle concepts based on hybrid wing body configurations have been proposed under NASA’s As previously mentioned, the distributed electric propulsion Subsonic Fixed Wing project to achieve low-noise and cruise- concept is not limited only to HWB aircraft but also could efficient short take-off and landing (CESTOL). The first easily be applied to other vehicle configurations such as vehicle concept was a high subsonic short take-off and landing traditional tube and wing aircraft and tilt rotor aircraft.

(STOL) capable hybrid wing body airframe with multiple, However, in order to achieve maximum benefits, it will be small, partially embedded conventional engines. The vehicle necessary to design an aircraft with greater emphasis on characteristics and performance data of that aerodynamically propulsion airframe integration right from the conceptual trimmed and low-noise concept vehicle were briefly reviewed design stage. Moreover, to achieve all the benefits described in in this paper. The present proposed vehicle is similar to the the above sections, a diligent research and development effort first but uses distributed superconducting electric fans, is required on the superconducting system for aircraft powered by two wing-tip-mounted turboelectric generators, to application. Besides additional modeling and analytical lower the fuel consumption, noise, and emissions even further, refinement of the electromagnetic, structural, and thermal as suggested by NASA’s SFW N+3 goals. Descriptions of the 7. Kim, H.D. and Saunders, J.D, “Embedded Wing vehicle, the superconducting system, and the propulsion Propulsion Conceptual Study,” NATO RTA Symposium th system were presented with some zero -order weight and on Vehicle Propulsion Integration, RTO–MP–AVT–100, efficiency comparisons to the multiple turbofan system. Oct. 2003.

Preliminary analysis suggests that fuel savings may be greater 8. Ko, A, Leifsson, L.T., Schetz, J.A., Mason, W.H., and than 6 percent for a turboelectric propulsion system compared Haftka, R.T., “MDO of a Blended-Wing-Body Transport to distributed discrete turbofans. Beyond fuel savings, Aircraft with Distributed Propulsion,” AIAA–2003–6732, however, turboelectric propulsion systems introduce a very Nov. 2003.

high degree of aircraft design and operational flexibility as a 9. Smith, A.M.O. and Roberts, H.E., “The Jet Airplane result of decoupling power production from power Utilizing Boundary Layer Air for Propulsion,” Journal of consumption. Lightweight superconducting generators, motors the Aeronautical Sciences , Vol. 14, No. 2, 1947, pp. 97– and power cables allow a small number of large turbo- 109.

generators to power an arbitrary number of propulsor units. 10. Küchemann, D., and Weber, J., Aerodynamics of Either can be placed practically anywhere and in various Propulsion , New York: McGrawHill, 1953.

orientations on the vehicle. This flexibility opens up design 11. Smith Jr., L.H., “Wake Ingestion Propulsion Benefit,” possibilities not obtainable with discrete turbofans or with Journal of Propulsion and Power , Vol. 9, No. 1, Jan-Feb.

distributed propulsion systems that employ mechanical power 1993.

distribution by gearboxes and shafts. 12. Spence, D.A., “The Lift Coefficient of a Thin, Jet- Flapped Wing,” Proceedings of the Royal Society of London. Series A, Mathematical and Physical Sciences,

VI. Acknowledgments

Vol. 238, No. 1212, pp. 46–68, Dec. 1956.

13. Williams, J., Butler, S.F., and Wood, M.N., “The The authors would like to acknowledge Ronald Kawai and Aerodynamics of Jet Flaps,” Aeronautical Research Sean Wakayama at Boeing Technology/Phantom Works, and Council Reports and Memoranda No. 3304, Jan. 1961.

James Stone and Eugene Krejsa at Diversitech, Inc., for their 14. Daggett, D., Hadaller, O., Hendricks, R., and Walther, R., contribution to the earlier cruise-efficient short take-off and “Alternative Fuels and Their Potential Impact on landing (CESTOL) vehicle analysis. We are also appreciative Aviation,” NASA/TM—2006-214365, Oct. 2006.

for the helpful interaction with Albert Kascak and Jeffrey 15. Guynn, M.D., Freeh, J.E., and Olson, E.D., “Evaluation of Berton at NASA Glenn Research Center, and Andrew Hahn at a Hydrogen Fuel Cell Powered Blended-Wing-Body NASA Langley Research Center for providing comments on Aircraft Concept for Reduced Noise and Emissions,” turboelectric CESTOL vehicle concepts. Finally, a special NASA/TM—2004-212989, Feb. 2004.

thanks goes to Fayette Collier, NASA’s Subsonic Fixed Wing 16. Kawai, R., “Quiet Cruise Efficient Short Take-Off and project principal investigator at Langley Research Center, for Landing Subsonic Transport System,” NASA/CR—2008- his support on current CESTOL activity.

215141, Apr. 2008.

17. Stone, J,R. and Krejsa, E.A., “Initial Noise Assessment of

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