Turboelectric Distributed Propulsion in a Hybrid Wing Body Aircraft

ISABE-2011-1340 Turboelectric Distributed Propulsion in a Hybrid Wing Body Aircraft James L. Felder, Gerald V. Brown, Hyun Dae Kim NASA Glenn Research Center Cleveland, Ohio 44135 USA Julio Chu NASA Langley Research Center Hampton, Virginia, 23681, USA bypass (UHB) geared turbofan engine with Abstract identical propulsion technology was integrated The performance of the N3-X, a 300 passenger into the same hybrid wing body airframe. That hybrid wing body (HWB) aircraft with aircraft achieved a 52% reduction in mission turboelectric distributed propulsion (TeDP), has fuel burn relative to the reference aircraft. The been analyzed to see if it can meet the 70% fuel N3-X was able to achieve a reduction of 70% burn reduction goal of the NASA Subsonic and 72% (depending on the cooling system) Fixed Wing project for N+3 generation aircraft.

relative to the reference aircraft. The additional The TeDP system utilizes superconducting 18% - 20% reduction in the mission fuel burn electric generators, motors and transmission can therefore be attributed to the additional lines to allow the power producing and thrust degrees of freedom in the propulsion system producing portions of the system to be widely configuration afforded by the TeDP system that separated. It also allows a small number of large eliminates nacelle and pylon drag, maximizes turboshaft engines to drive any number of boundary layer ingestion (BLI) to reduce inlet propulsors. On the N3-X these new degrees of drag on the propulsion system, and reduces the freedom were used to (1) place two large wake drag of the vehicle.

turboshaft engines driving generators in freestream conditions to maximize thermal Nomenclature efficiency and (2) to embed a broad continuous A area array of 15 motor driven propulsors on the upper AC alternating current surface of the aircraft near the trailing edge. That ADP aerodynamic design point location maximizes the amount of the boundary BLI boundary layer ingestion layer ingested and thus maximizes propulsive BSSCO bismuth strontium calcium copper efficiency. The Boeing B777-200LR flying 7500 oxide nm (13890 km) with a cruise speed of Mach CAEP Committee on Aviation 0.84 and an 118100 lb payload was selected as Environmental Protection the reference aircraft and mission for this study.

C nozzle throat discharge coefficient dth In order to distinguish between improvements C nozzle velocity coefficient V due to technology and aircraft configuration DC direct current changes from those due to the propulsion FPR Fan Pressure Ratio configuration changes, an intermediate HWB hybrid-wing-body configuration was included in this study. In this ISA International Standard Atmosphere configuration a pylon mounted, ultra high LTO landing and take-off Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

K Kelvins General Electric (GE) . Northrop Grumman, M Mach number Boeing and MIT examined medium size and MgB magnesium di-boride range aircraft in the 150 passenger class. MIT MIT Massachusetts Institute of also examined a large intercontinental 300 Technology passenger hybrid wing body (HWB) aircraft.

NRA NASA Research Announcements The team led by GE elected to examine short NOx oxides of nitrogen range 20 passenger aircraft flying point to point nm nautical mile between the thousands of smaller airports P pressure distributed broadly around the United States.

R degrees Rankine A NASA internal team composed of the authors RTO rolling take-off felt that a radical departure in both aircraft and TeDP turboelectric distributed propulsion propulsion system was needed to meet the N+3 UHB ulta-high bypass goals. An intercontinental mission of 7500 nm YBCO yttrium barium copper oxide (13890 km) with a cruise speed of Mach 0.84 and a 118100 lb payload was selected as the Introduction reference mission for this study. The 300 The NASA Subsonic Fixed Wing (SFW) project passenger Boeing B777-200LR was selected as has defined goals for the next three generations the reference aircraft against which to compare of aircraft in four key areas of reducing noise, mission fuel burn. A reference model patterned fuel burn, emissions and field length. Table after the B777-200LR was constructed using the 1 outlines goals for each generation. The dates NASA Flight Optimization System (FLOPS) given for each generation are targets for code . The predicted fuel burn from FLOPS for attaining technology readiness levels (TRL) 4 to the reference aircraft/engine combination flying 6 . The NASA SFW project formed six teams, the reference mission is the value against which two internal and four external, to examine candidate aircraft/engine mission fuel burns are concepts to meet the N+3 goals. The external compared.

teams selected during the NASA Research The B777-200LR is powered by the GE90-115B Announcement (NRA) phase 1 study were led engine. A reference model of an engine by the Massachusetts Institute of Technology patterned after the GE90-115B was constructed 2 3 4 (MIT) , Northrop Grumman , Boeing and using the Numerical Propulsion System N+1 (2015)*** N+2 (2020)*** N+3 (2025)*** Technology Benefits Technology Benefits Technology Benefits CORNERS OF THE Relative to a Relative to a TRADE SPACE Single Aisle Reference Large Twin Aisle Reference Configuration Configuration Noise - 32 dB - 42 dB - 71 dB (cum below Stage 4) LTO NOx Emissions -60% -75% better than -75% (below CAEP 6) Performance -33%** -50%** better than -70% Aircraft Fuel Burn Performance -33% -50% exploit metroplex* concepts Field Length *** Technology Readiness Level for key technologies = 4-6 ** Additional gains may be possible through operational improvements * Concepts that enable optimal use of runways at multiple airports within the metropolitan areas Table 1 NASA Subsonic Fixed Wing Goals for the Next Three Aircraft Generations Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

Simulation (NPSS) code . The NPSS model is Ideally, this would be done without labeled as the Pax300. causing the nozzles to project beyond the trailing edge for noise reduction An HWB configuration was selected for the new reasons.

aircraft. The HWB type aircraft presents an opportunity to reduce both fuel burn and aircraft  Continuous inlets and nozzles to noise. HWB aircraft present a relatively good lift minimize external wetted area and avoid to drag ratio (L/D) of around 22 , leading to additional drag due to channel flow reduced fuel burn. The engines can be mounted between closely spaced nacelles.

on the upper aft portion of the center body where  Core engines that do not ingest the the fuselage can potentially provide noise 9 boundary layer in order to avoid losing shielding . However, placing the engines on top thermal efficiency.

of the fuselage presents a number of challenges.

 A minimum number of core engines to For pylon mounted turbofans these challenges maximize thermal efficiency due to include a high thrust line relative to the aircraft larger turbomachinery. The ideal would center of gravity. Another challenge is either a be only two core engines which provide high inlet Mach number (if the engines are the minimum for engine-out sufficiently far forward to provide fuselage redundancy.

shielding of the exhaust noise) or loss of fuselage noise shielding (if the engines are  A power distribution method with high moved aft to avoid the high velocity portions of transmission efficiency.

the center wing-body). Embedding the engines A number of recent studies [ 2 , 10 , 12 , 13 ] have into the upper aircraft surfaces addresses the examined the use of single-fan and multi-fan high thrust center while also eliminating the turbine engines embedded in the upper surface weight and drag of the pylon and a portion of the of an HWB aircraft. The predicted fuel burn nacelle. Embedded engines can also reductions due to BLI have been in the 3%-7% significantly reduce fuel burn by ingesting the range compared to a pylon mounted engine of boundary layer. Ingesting the boundary layer equal technology level. Single fan turbofans reduces the average inlet velocity to less than the require heavy, high-aspect-ratio inlets if a freestream value and thus reduces the drag of the significant portion of the boundary layer is to be inlet. If the inlets can also be located far aft on ingested and if the number of engines is limited the HWB center body airfoil section, the natural to two or three. Performance losses due to diffusion of the airfoil will also reduce the additional internal pressure loss and additional velocity of air above the boundary layer. This inlet distortion can more than off-set the gains further reduces inlet drag for inlets that project due to BLI. Low aspect ratio inlets avoid the above the boundary layer height. However, losses of high aspect ratio inlets, but also limit ingesting the boundary layer can result in the amount of boundary layer ingested, resulting significant losses. As documented by Tillman , in only small improvements in fuel burn.

it is easy for the losses associated with boundary Increasing the number of engines allows the use layer ingestion (BLI) to more than off-set the of more low-aspect ratio inlet to ingest the same gains.

percentage of the boundary layer. However, To maximize gains while minimizing losses, the smaller engines may be limited to lower overall optimal BLI propulsion system on a HWB pressure ratios (OPR) than larger engines and aircraft would have the following attributes: they are more susceptible to adverse effects like tip clearance and surface finish, all of which  Inlets that ingest a large percentage of reduce thermal efficiency.

the upper surface boundary layer.

The multi-fan approach where a single core  Inlets located near the trailing edge to engine drives multiple fans either through take full advantage of BLI and aft airfoil mechanical, hydraulic or hot gas power diffusion to reduce inlet velocity.

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Figu re 1 N3 - X Hybrid Wing Body (HWB) Aircraft with a Turboelectric Distributed Propulsion (TeDP) distribution addresses some of the issues seen devices to be independent of each other, with single-fan configurations, while adding essentially forming a infinitely variable ratio some of its own. The larger number of smaller transmission between the power turbines and the fans allows more of the boundary layer to be fans. Also electrical power from multiple ingested by low aspect ratio inlets while devices can be readily mixed, allowing a degree maintaining the thermal efficiency of a few of cross connection that is very difficult to larger core engines. This approach, however achieve with mechanical power distribution.

adds the weight and losses of a right-angle drive Where other embedded engine concepts meet gearbox, hot gas ducting or hydraulic pumps and some of the criteria, we feel our design meets all motors. The predicted results of the multi-fan of the criteria outlined above for an optimum approach have ranged from a small decrease to a BLI system.

small increase in fuel burn relative to the Configuration and Assumptions standard pylon mounted engine of equal technology level.

The turboelectric distributed propulsion (TeDP) system illustrated in Figure 1 consists of two The authors of this paper elected to examine the turbogenerators consisting of a turboshaft engine use of electrical energy to transmit power from the gas turbines to the fans. Transmitting all driving superconducting electrical generator.

The primary function of these devices is to make power between the turbines and the fans as electricity, not thrust. The nozzle of the electricity allows the power generator and the propulsors to be placed anywhere on the vehicle turbogenerator is sized so that there is enough jet velocity at cruise to produce a small amount of to optimize overall system performance.

net thrust to avoid being a source of drag. They Electrical power can be transmitted long distances with very little loss. The flexibility in are located on the wingtips so that the inlets ingest freestream air. Most of the energy of the distributing electrical power allows the number gas stream is extracted by the power turbine to of power producing devices and the number of thrust producing devices to be independent of drive the generator. As a result the exhaust velocity is low which should result in low jet one another. Distributing the power as direct current (DC) allows the speeds in the different noise as well. The wingtip location will also give Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

some bending moment relief in the normal will fit across the given array width with at least direction at the cost of an increase in bump a minimum specified separation between fans.

loading and possible aeroelasticity The requirement that the propulsor array be considerations.

continuous sets the width of each individual There are other potential benefits of locating the propulsor inlet to equal the total width of the turbogenerators on the wingtips. Research array divided by the number of propulsors. The conducted in 1970 at NASA identified inlet height is determined by using the reductions in induced drag of up to 40% if a relationship between mass-averaged Mach thrust producing device is located at the wing number versus height above the fuselage for the tip. This reduction is due the higher velocity specific aircraft shape and percent chord thrust stream reducing the strength of the wing- location to determine the inlet height. The mass- tip vortex well downstream of the wing itself. average Mach number and density determine the While the aspect ratio of the wing and the flow mass flow into an inlet of the given height. The rate of the turbogenerators is different than the height of the inlet is iterated and the mass- configuration tested in the wind tunnel, the average Mach number reevaluated at each height basics of the configuration are the same and so until the height is such that the inlet mass flow there should be some induced drag reduction equal to the fan mass flow. This iterative Another argument for the wingtip location is that calculation is done as part of the design point it nearly eliminates the risk to the rest of the calculations in the NPSS model of the TeDP aircraft and passengers in the event of a turbine system. And lastly the nozzle height is disk burst. Future analysis will further quantify determined from the calculated nozzle area these effects. The wingtip location is not divided by the array width.

mandatory. The turbogenerators can be Each propulsor in the array consists only of a embedded in the wing root area with a leading low aspect ratio two-dimensional (2-D) slot edge inlet or on short pylons on top of the wing inlet, a fan, a short duct around the motor and a if needed without sacrificing high inlet pressure low aspect ratio 2-D slot nozzle. The result is a recovery or incurring large installation losses.

very short axial length for each propulsor. This The electric power from the turbogenerators is allows the inlets to be located further aft to distributed along redundant superconducting maximize BLI benefits while still allowing the electrical cables to an array of superconducting nozzle to be located forward of the trailing edge.

motor driven fans in a continuous array of Thus fuselage noise shielding of the propulsor propulsors spanning the entire upper trailing stream is maintained. This also means that the edge of the center wing-body section. The width propulsion system does not cover the pitch of the array is set to cover all of the long chord effector and thus requiring thrust vectoring of portions of the fuselage and wing root. This the engines to control aircraft pitch.

maximizes the amount of boundary layer An HWB aircraft derived from the Boeing ingested as measured by the swept area ahead of 12 13 N2A& N2B and SAI SAX-40 with the the propulsors with a minimum number of addition of aircraft technologies anticipated to propulsors.

be available in the N+3 timeframe was used in The number of propulsors is not set, but rather is this study. A model of this aircraft was a function of the width of the array, the fan constructed in FLOPS. Two versions of this pressure ratio (FPR) and the net thrust that is aircraft were created. The TeDP system is required. For a given FPR, 1.3 for this study, integrated into the first version. The resulting and a given amount of thrust required by the TeDP/HWB combination, seen in Figure 2 , is aircraft the aggregate area of all the fans can be referred to as the N3-X. The second version has calculated. The number of fans, and thus two ultra high bypass (UHB) geared turbofans propulsors, is then determined by the number of mounted on pylons on the upper surface. The circular fans of the required aggregate area that UHB turbofan is assumed to have the same component efficiencies and material temperature Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

limits as assumed for the TeDP. The UHB/HWB is very different gross take-off weights and thus configuration, seen in Figure 3 , is very similar to thrust levels for N3 aircraft with the two engine the Boeing/NASA N2A and so is referred to as types.

the N3A. Engine simulations of the TeDP and Table 2 gives the uninstalled thrust required for the UHB were constructed using the NPSS the N3A and N3-X. These thrust values were used to size the two propulsion systems.

The nature of the TeDP configuration is such that there is very little installation drag. This is because the only extra external wetted area over the basic airframe is the sides of the propulsor array and the turbogenerator nacelle. The top of the propulsor nacelle has the same wetted area as the aircraft fuselage section that it covers.

Thus no installation drag penalties were assessed against the TeDP on the N3-X configuration.

The UHB, however, does have installation drag Figure 2 N3 - X FLOPS Model Configuration Flight Uninstalled Condition Thrust lbf(kN) program. Engine performance from the TeDP N3A RTO 78766 (350.37) and UHB simulations are then input into the ADP 25378 (112.89) FLOPS models of the vehicles to allow vehicle N3 - X RTO 54888 (244.15) sizing and mission analysis to be performed.

ADP 19293 (85.82) Table 2 N3 A and N3 - X Uninstalled Thrust Requirements associated with the nacelle and pylon. The result is that comparing uninstalled performance of the two engine types would be misleading. Thus performance of the two engine configurations can be compared only on an installed basis.

Boundary Layer Conditions A detailed understanding of the inlet flow field is critical to correctly estimating the Figure 3 N3A FLOPS Model performance of a BLI propulsion system as demonstrated by the authors in prior work . The N3-X and N3A Thrust Requirements results of a three-dimensional (3-D) computational fluid dynamics (CFD) simulation The aircraft thrust requirements for the N3-X of the closely related N2A-EXTE aircraft by and N3A are defined at two flight conditions; rolling take-off (RTO) at sea level, Mach 0.25, Friedman was used to estimate the boundary layer Mach number and total pressure profiles at ISA+27R hot day, and the aerodynamic design a range of percent chord locations along the point (ADP) at 30000 ft, Mach 0.84, ISA centerline. The N2A-EXTE represents an standard day. Even with the same technology extension of the tail region by about 200 inches assumptions for the turbomachinery, the to a chord length of 1800. This was done to differences in configuration between the N3-X provide additional aft fuselage for noise and N3A engines yield a considerable difference shielding. The N3-X does not include this in installed specific fuel consumption. The result Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

along the centerline of the N2A-EXTE. Sizing 1.05 the propulsors around a 1.3 fan pressure ratio 1200 1.00 resulted in the inlet plane being located at 1000 0.95 the 85% chord location.

Therefore the boundary 0.90 layer profile for this location was used to 0.85 represent the boundary layer entering the 0.80 propulsor inlets.

0.75 The velocity and density of the flow at each 0.70 location in the boundary layer at the 85% chord 1800 1600 1400 1200 1000 800 600 400 200 0.65 location was used to calculate the mass flow per unit area at all values of height in the Figure 4 N2A - EXTE Upper Surfac e Isentropic Mach Number boundary layer. These Distribution mass flow rates were used to determine the mass-averaged Mach number and total 30 pressure for a given inlet 25 x/c 0.6 height. This was x/c 0.7 repeated for all heights x/c 0.8 to give curves of mass- x/c 0.85 average Mach number inches - x/c 0.9 and total pressure versus x/c 1.0 inlet height. To extend height the usage of these curves beyond the flight condition of the original 0.00 0.20 0.40 0.60 0.80 1.00 CFD simulation, the curves were normalized Mach number by the freestream Mach number and total Figure 5 N2A - EXTE Centerline Boundary Layer Mach Number pressure at which the Profiles CFD was run. The extension. However, the differences are small resulting PtRatio and MNratio are given in enough that it was judged that the boundary Figure 6 .

layer shape and height would be the same on the These two curves are central to estimating the two aircraft at the same percent chord locations.

effect of ingesting the boundary layer. When Figure 4 shows the inviscid Mach numbers at sizing the propulsors at the ADP flight condition the top of the boundary layer for the top of the the freestream Mach number and total pressure aircraft. Figure 5 shows the Mach number of 0.84 and 6.93 psia are used to unnormalize profiles for a range of percent chord locations Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

inches - PtRatio 30 MNratio height 0.6 0.7 0.8 0.9 1 Fraction of Freestream Value Figure 6 x/c = 0.85 Mass - avg PtRatio & MNratio the curves. An estimate of the required inlet the inlet height due to the continued decrease in height is made and the mass-average Mach average velocity with reduced capture height.

number for that height is used to determine the When the capture height is higher, then there is inlet mass flow. Also with the mass-averaged external acceleration to contract the flow down Mach number and total pressure the diffusion or to the physical inlet height. The external ram drag resulting from the inlet is also acceleration causes a drop in the external static calculated. With the inlet mass flow and total pressure. This suction effect along the trailing pressure the thrust produced by each propulsor edge will enhance the circulation around the can be calculated. This is compared to the fuselage leading to an increase in the lift required thrust, and if not equal then the inlet coefficient. Quantifying the change in the lift height is varied until they are.

coefficient due to inlet suction and its impact on A similar process is used in off-design balance field length will be the subject of future calculations to determine the capture height of analysis.

the inlet stream (assuming that the width of the The wing-tip location results in freestream inlet inlet stream does not vary). If the capture height conditions to the inlets of the turbogenerators, is less than the inlet height, then the mass- and so the propulsor inlet conditions have no average inlet Mach number and total pressure effect on the turbogenerator performance.

decreases. Also when the capture height is less Turbomachinery Design than the inlet height there is external diffusion sufficient to expand the flow from the capture The N+3 performance goals represent an height to the inlet height. Effects of the external extreme technical challenge. To assess the diffusion on the flow field in front of the inlet ability of the TeDP/HWB concept to reach the are not captured at this time.

goal of 70% reduction in mission fuel burn, an optimistic approach to estimating the design The benefit from BLI is captured in the reduced parameters was taken to determine if the fuel ram drag that results from the mass-average inlet burn goal was obtainable even with optimistic velocity being lower than freestream velocity.

assumptions. All turbomachinery efficiencies, The BLI benefit is higher at conditions, such as temperature limits and material assumptions part power, where the capture height is less than Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

Componen Parameter Design Value/Assumptions t Inlet Geometry 2 - D “mailslot”. Width equal to fan diameter plus spacing between adjacent fans. Height calculated from flow area divided by width.

dP/P (throat to fan) 0.002 Fan PR 1.30 Adiabatic efficiency 0.9535 Distortion efficiency penalty 0.01 Design Tip Speed 883 ft/sec Nozzle Geometry 2 - D low aspect ratio with variable exit area Cdth 0.997 Cv 0.997 Table 3 Propulsor Design Parameters Component Parameter Design Value/Assumptions LPC & HPC Polytropic efficienc y 0.9325 HPC Maximum exit total 1810 R @ RTO, 1681 R @ ADP temperature (T3) LPC & HPC Pressure Ratio (PR) Total PR varied to equal max T3 with an equal ∆h (enthalpy) split between compressors Burner Exit total temperature (T4) 3360 R @ RTO, 3260 R @ ADP HPT & LPT Polytropic efficiency 0.93 PT Polytropic efficiency 0.924 Turbine Ceramic Matrix Com posite Uncooled for all hot section components including material (CMC) burner liner, and turbine stators and rotors with 3460 R max material temperature HPT Non - chargeable disk cooling 4% LPT Non - chargeable disk cooling 2% PT chargeable disk cooling & 1% cavity p urge Nozzle PRdes 1.6 @ 30k/MN0.84 ADP PRmin 1.01 Table 4 Turboshaft Engine Design Parameters used in the TeDP system were applied to UHB impacts of BLI can be limited to efficiency geared turbofan engine as well. penalties of 1% to 2%. The baseline fan efficiency represents N+1 technology, so we The design assumptions for the propulsor are assessed only a 1% efficiency penalty. The same given in Table 3. The fan efficiency and design optimization that reduced the impact on tip speed were first obtained from a NASA N+1 efficiency also reduced the internal total study of engines for single-aisle transports.

pressure loss of the embedded inlet to 0.2%- Analysis by Tillman indicates that with 0.3%. The relatively low fan pressure ratio optimization of the inlet geometry the distortion necessitates a variable area propulsor nozzle.

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The segmented 2- D “mail - slot” nozzle allows current motor and generator power levels of the upper surface of the nozzle to be simply 4000 hp and 30000 hp respectively. The hinged to provide the necessary variability. resulting weights and efficiencies of each of the electrical components is presented in Table 4 contains the design efficiencies and Table 7 for magnesium di-boride (MgB ) and in temperatures for the turbogenerator. The Table 8 for bismuth strontium calcium copper pressure ratio split between the low pressure and oxide (BSCCO). The power levels used to size high pressure compressors was varied such that the electric system are those needed at the there was an equal enthalpy rise across each rolling take off (RTO) flight condition. This is compressor. NASA materials roadmaps for done because the electrical portion of the system ceramic matrix composites (CMC) anticipate a must be sized to handle the highest power maximum material temperature of 3460R. With transmitted, which is at take-off. . The this material temperature limit, turbine blade turbomachinery portions of the propulsion cooling is unnecessary. The burner exit pressure system are sized at the aerodynamic design point was set at 3360R to give a 100R factor of safety.

(ADP).

Some cooling flow is still required to cool the turbine rotor disks and for cavity purge. The Fully Superconducting Generator nozzle pressure ratio was set to yield minimum thrust at cruise. The result is that more of the energy in the gas stream is extracted by the power turbine and directed to the generator and less is left in the exhaust flow. The result is a lower exhaust gas velocity from the turbogenerator, especially at the RTO condition where noise is critical. If the turbogenerator noise is found to be an issue in meeting the N+3 noise goal a possible option is a variable area turbogenerator exhaust nozzle. This would allow Figure 7 Schematic Drawing of a Fully the nozzle area to be increased at high power Superconducting Electric Machine settings to further increase the pressure ratio across the power turbine and decrease the nozzle exhaust further from what is possible with a This study assumes that the required power fixed area nozzle. At cruise conditions the density of the motors and generators is obtained nozzle area would be reduced to produce an from wound rotor synchronous machines with optimal amount of thrust directly from the superconductor windings on both rotor and turbogenerators.

stator. The state-of-the-art of superconducting machines has been reviewed in various papers Electrical Power System cited in reference 17 . Such a machine is In the following sections the components depicted schematically in Figure 7 . The stator required in a superconducting electrical power windings deliver AC power to be rectified and system are listed and briefly discussed. The transmitted as DC power by superconducting methods underlying this section are the same as transmission lines. The high power electrical presented in Brown’s paper on weights and components all operate at cryogenic efficiencies of electric components of a temperatures without any electrical leads turboelectric aircraft propulsion system . That between cryogenic temperatures and room same analysis has been repeated here for the temperature.

Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

The windings in the stator of the generator are can be high enough to reduce the power turbine subject to alternating magnetic fields and weight substantially below the value required for alternating currents (AC), and therefore suffer a direct drive large fan. Thus electric drive superconducting AC losses. To keep the provides the same advantage with respect to required cooling capacity within acceptable turbine weight as a planetary gear box.

bounds, these stator conductors must be carefully engineered to reduce the AC losses.

This can be achieved by using small diameter wire with fine superconducting filaments embedded in a resistive metallic matrix. The required filament size presently appears achievable only for the superconductor MgB . A 2009 NASA Small Business Innovation Research (SBIR) contract conducted by Hyper Tech Research, Inc examined the production of MgB suitable for use in turbo-electric aircraft propulsion system. Figure 8 shows the cross section of their interim design. Unfortunately, the critical temperature (the highest temperature at which it is superconducting) for MgB is only 39K, and it must operate below 30K to yield a useful current density. This increases the weight of the required cryocooler. Even though BSCCO cannot currently be formed with acceptable AC losses, a future development is assumed that will make it possible. Motors and generators with a Figure 9 Turbo - Brayton Cryocooler hypothetical fine-wire BSSCO are included in this study. The superconductor yttrium barium Cryocoolers A cryocooler is a refrigerator that produces very low temperatures. The superconducting devices used in the TeDP system require temperatures between 20K and 65K. A 2009 NASA SBIR study conducted by Creare produced a preliminary design of a turbo-Brayton cryocooler, depicted in Figure 9 , which meets our current weight goal of 5 lb/hp-input. This is th about 1/6 the weight of the best present coolers and is expected to achieve the same 30% of Figure 8 Low - AC Loss Superconducting Carnot efficiency attained by the best current Configuration technology coolers. A more appropriate functional dependence of the weight on the input power (instead of simple proportionality) is not copper oxide (YBCO) was not considered for yet within our modeling capability.

this application because no concept for suitably The cryocoolers are driven by their own electric low AC loss has yet been advanced for this motors, which are included in the weight material.

estimate. The power to drive the cryocoolers Note that the generator shaft speed can be comes from the turbogenerators. Thus the power chosen to match the optimum speed of the losses in the superconducting devices and the power turbine in the turbine engine. that speed inverters plus the power to the cryocoolers Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

represent the total transmission power loss. The combination of latent heat and sensible heat amount of power required by the cryocoolers capacity.

depends on the operating temperature of the After serving as a coolant, the hydrogen, with a device being cooled (the source temperature) lower heating value of 51585 BTU/lb, is and the temperature at which heat is being compressed and introduced to the burner of the rejected (the sink temperature). The larger the turbogenerator where it replaces a portion of the difference between the source and sink jet fuel equal to approximately 2.8 times the temperatures, the greater the cryocooler power mass of the hydrogen. The amount of hydrogen required. It is for this reason that the cryocooler required is very small compared to the energy power and weight is higher for MgB based needed to propel the aircraft and so jet fuel still devices which operate below 30K than for constitutes the majority of the fuel energy.

BSCCO based devices which operate near 50K.

Superconducting Transmission Lines Liquid Hydrogen Cooling The superconducting transmission line has not Liquid hydrogen was examined as an alternate to yet been studied in detail for this application.

cryocoolers to cool the superconducting motors, Many studies and demonstration projects for generators and transmission lines as well as the ground transmission lines for utility grids have cryogenic inverters. Liquid nitrogen does not been made or are under way. Either AC or DC present an alternative. This is because even transmission is possible. A 2006 Chinese test of though the BSCCO material has a critical a superconducting, 60 Hz AC cable was temperature above the boiling point of liquid reported. The 3-phase cable carried 120MW nitrogen, the critical current density at liquid (over twice the N3-X take-off power of 45 MW) nitrogen temperature is too low to yield motors, and the mass of each phase was 9.2 kg/m .

generators and transmission line with weights Losses are only a few watts per meter.

useful in an aircraft application. The critical Superconducting cables typically operate at current density rises sharply with decreasing liquid nitrogen temperature, 77K. But cooling to temperature. With a boiling point of 20.4K at approximately 55K, to match the motors and ambient pressure, liquid hydrogen provides an generators of the TeDP system, would increase operating temperature that yields very high the critical current density and allow about 3 current densities, resulting in smaller and lighter times the power capacity for the same cable size.

motors, generators and power lines. Liquid Pending detailed studies, a weight of 1000 lb for hydrogen can also directly cool MgB based superconducting cables was added to the total machines, which need to operate at 30K or less.

electrical system weight. Transmission losses In the refrigerated system, cryocoolers represent and environmental heat transfer combined are the largest power “loss” in the transfer of power typically on the order of 5 W/m of cable length.

from the power turbines to the fans. The use of In addition to its central task of carrying current, liquid hydrogen eliminates this particular loss the transmission lines, which are built around a and substantially increases the power hollow core to carry a coolant for its own transmission efficiency. The hydrogen cooling cooling, can also be used to carry coolant to flow rate required is calculated by assuming that motors or generators, avoiding a separate hydrogen boils at constant temperature in the coolant line and allowing a central location for AC stator with the cold gas that is evolved the cryocoolers or hydrogen tanks.

cooling the nearly lossless DC rotor. The Fully Superconducting Motors hydrogen gas from the motors then travels to the associated inverters to cool those devices. If the An electric machine can operate as either a motor or a generator. The superconducting heat capacity of the hydrogen flow required to motors were sized with the same sizing code and cool the motor is not sufficient to cool the inverter, then additional liquid hydrogen is treated in exactly the same fashion as the generators. It was assumed that they are driven introduced to remove the remaining heat with a by cryogenically cooled inverters so that the Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

shaft speed of the fans can be varied power of 15 hp/lb, including the cryocooler .

independently of the generators. For the example This specific power is used here to estimate the aircraft, N3-X, of Fig. 1, fifteen motors of 4,000 weight of the inverters for the given power level.

hp each are required. With a maximum fan tip At cruise power levels the inverter efficiency speed of 883 ft/s (for a 1.3 FPR) and a calculated could be as high as 99.93% for the inverter fan diameter of 43 inches, the fan motor shaft itself. Including power to drive a cryocooler to speed is calculated to be 4500 rpm. provide the necessary temperature the combined efficiency is still 99.5%. The efficiency at Cryogenic Inverters takeoff power levels would be about 0.25% Inverters convert direct current to alternating lower, but the duration of takeoff is short, so the current of any desired frequency and therefore impact on total fuel burn is not significant.

can drive a motor at a variable speed. Inverters allow the fan speed to be independent of the engine’s power shaft speed, in effect acting as a Analysis and Results variable ratio gear box. This is a key factor in allowing all propulsors to continue to operate in Engine Performance Table 5 and Table 6 compare the performance of the TeDP the event that a turbogenerator fails. In that system to the performance of the geared UHB situation the propulsor speeds would drop to the point where the power demand equals the power turbofan and the NASA developed Pax300 direct drive turbofan . The Pax300 model is from the remaining turbogenerator. The similar to the GE90-115B present on the remaining turbogenerator, meanwhile, would reference aircraft, the Boeing B777-200LR. It is included here to give a current technology metric. The UHB and TeDP engines were 1.4 iterated with the FLOPS models of the N3A and 1.2 N3-X aircraft to determine the fuel load UHB necessary to perform the reference mission and 1.0 thus determine both the empty and gross take-off TeDP vehicle weights. The low fuel consumption of 0.8 the TeDP resulted in a smaller fuel load and thus 0.6 a smaller vehicle that required less thrust that in turn reduced the size of the TeDP engine.

0.4 Both the UHB and TeDP engines are sized to a 0.2 Installed TSFC (lbm/hr/lbf) fan pressure ratio (FPR) of 1.3 at the 0.0 aerodynamic design point (ADP) and a T3 of 1810R and T4 of 3360 R at the rolling take-off 0.0% 20.0% 40.0% 60.0% 80.0% 100.0% (RTO) flight conditions while meeting the thrust Percent Max ADP Thrust required at both flight conditions. The thrust lapse rates with altitude and speed for both the UHB and the TeDP cycles were such that Figure 10 UHB & TeDP Installed TSFC Vs Percent of engines were designed to provide more thrust at Maximum Thrust at ADP the ADP point than the aircraft required in order have to maintain or even increase speed and to match the thrust required at the RTO flight power output. Various factors related to the condition. The UHB is capable of 5% more inverters are discussed in reference 17 , including thrust at the ADP flight condition than the N3A the higher efficiency and lower weight of requires, while the TeDP system is capable of cryogenically cooled inverters. These 20% more thrust.

advantages hold even when the required cryocooler weight and efficiency is taken into Requiring the TeDP system to be oversized by account. A 2010 NASA SBIR has predicted that 17% at the ADP point would ordinarily cause a cryogenic inverter could attain a specific Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

the cruise fuel efficiency to be RTO Reference N3A TeDP TeD P penalized since standard engine Engine ( Cryo ) ( LH2 ) behavior is for the TSFC to Altitude (ft) 0 0 0 0 increase with decreasing thrust Mach number 0.25 0.25 0.25 0.25 for a given flight condition.

However, the TeDP does not dTs ( R ) 27 27 27 27 exhibit this behavior. Figure 10 Pt (ambient) 15.35 15.35 15.35 15.35 compares the TSFC versus percent power trends for the MN (capture) 0.25 0.25 0.233 0.233 UHB and TeDP engines. The Pt (capture) 15.35 15.35 14.94 14.94 UHB engine exhibits the T3 ( R ) 1673 1803 1791 1789 standard “power - hook” trend of small increase in TSFC from T4 ( R ) 3296 3360 3358.4 3356 100% to 80% thrust with ever Fn (Installed) 161215 78249 54888 54882 faster increase in TSFC with further throttling. The TeDP 63134 20177 13807 11867 W (lb/hr) fuel engine displays a completely different trend. For the TeDP 0.3919 0.2578 0.2515 0.2162 TSFC (installed) the TSFC declines in a (lbm/hr/lbf) continuous manner such that W (lb/s) 6503 6539 4940 4944 air the lowest TSFC is actually at BPR 8.8 29.5 32.7 34.2 idle. At 83% of maximum thrust, the TSFC of the TeDP is FPR 1.4 9 1.2 1.2 1.2 5% lower than at maximum OPR 43.1 57.4 56.1 55.8 thrust. The reason for this 286.3 286.3 286.3 286.3 V (f t/s) unusual trend is that the effect amb of BLI increases at part power.

286.3 286.3 267.1 267.1 V (ft/s) capture Table 7 and Table 8 present 923 665 618.6 618.6 V (ft/s) bypass weights and efficiencies of the 1126 898 817 818 V (ft/s) core major electrical components of the TeDP system on the N3-X 47.3% 60.2% 64.6% 64.6% η (bypass) Propulsive aircraft. A generator with its 40.5% 48.3% 52.8% 52.8% η (core) Propulsive cooler was optimized with 46.7% 59.8% 64.3% 64.3% respect to several design η (avg) Propulsive parameters to minimize the Fan Diameter (in) 128 150 41 41 combined weight. A motor 25736 35249 19746 19784 A (vehicle) (in2) fan with its cooler was separately optimized in the same way.

The total electrical system Table 5 Rolling Tak e - off (RTO) Engine Performance weights are for a system with weights of the cryocoolers, but does not include 15 propulsors and 2 turbogenerators. Each an estimate of the hydrogen tankage weight, propulsor has one motor and one inverter. Each which dependents on hydrogen required for the turbogenerator has one generator. The design mission. The weight of the hydrogen transmission line weight represents the estimate coolant itself is included in the fuel weight of transmission line weight for the entire rather than the electrical system weight.

aircraft. The weight for the cyrocooled system includes the weights of the motor, inverter and The operating temperature for MgB based generator cryocoolers. The weight of the LH2 devices of less than 30K results in higher cooled system in these tables excludes the cryocooler power and larger, heavier Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

hydrogen contained is being ADP Reference N3A TeDP TeDP explored. About 2370 lbs of Engine ( Cryo ) ( LH2 ) liquid hydrogen is required to 30000 30000 30000 30000 Altitude (ft) cool the MgB system for the 0.84 0.84 0.84 0.84 M (amb) reference mission. This would require a hydrogen tank 0 0 0 0 dTs ( R ) weighing about 1185 lbs, for a 6.93 6.93 6.93 6.93 Pt (ambient) net empty weight reduction of 4861 lbs, compared to the 0.84 0.84 0.735 0.735 M (capture) cryocooled MgB .

6.93 6 .93 6.48 6.48 Pt (capture) Second, the hydrogen can be 1510 1670 1603 1603 T3 ( R ) used as fuel after it is used as 3029 3212 3049 3051 a coolant. Hydrogen has a T4 ( R ) lower heating value (LHV) of 55697 24173 19293 19293 Fn (Installed) 518585 BTU/lb while jet fuel 31495 11281 6659 5673 W (lb/hr) fuel has a LHV of 18580 BTU/lb.

Thus one pound of hydrogen 0.5780 0.4667 0.3451 0.294 TSFC (installed) can replace about 2.8 pounds (lbm/hr/lbf) of jet fuel. Thus the 2370 lbs W (lb/s) 3501 3485 2503 2505 air of hydrogen replaces 6615 lbs BPR 8.5 27.2 30.5 31.7 of jet fuel for a net benefit of 4245 lb reduction in total fuel FPR 1.587 1.290 1.26 1.26 Third, the efficiency of OPR 43.1 71.1 64.1 64 transferring power from the 835.8 835.8 835.8 835.8 V (ft/s) amb engines to the fans rises from 835.8 835.8 742.3 742.5 V (ft/s) 97.75% to 99.88% without capture power being consumed by the 1040 1006 986.8 986.8 V (ft/s) bypass cryocoolers. This 2.1% 1587 1418 1164 1169 V (ft/s) core increase in transmission efficiency will translate 89.1% 90.8% 96.7% 96.7% η (bypass) Propulsive directly into reduction in total 69.0% 74.2% 87.7% 87.4% η (core) Propulsive fuel weight.

87.0% 90.2% 96.4% 96.4% η (avg) Propulsive Fourth, the necessary hydrogen can be generated from non-carbon emitting Table 6 Aerodynamic Design Point (ADP) Engine Performance sources of power. Also cryocoolers than are required for BSCCO based hydrogen generators located at the airport can systems which operate at 50K or more. With a serve as load of last resort for renewable power boiling point at atmospheric pressure of 20.4K, sources like wind, solar, or wave for times when liquid hydrogen represents an attractive more power is being generated than can alternative to cryocoolers for MgB based otherwise be used. Coordination between devices. The advantages of hydrogen are airports and local electrical companies can use fourfold.

the hydrogen generation systems to help balance the load on the entire electrical grid in a way that First, the 6064 lbs weight of the cryocoolers provides benefits to both parties.

listed in Table 7 is eliminated. Partially offsetting this is The impact of change in efficiency plus the the need to carry cryogenic tanks to hold the effect of higher LHV reduces the fuel weight a hydrogen. Technology for hydrogen tanks that total of 6050 lbs compared to an MgB system are one third to one half the weight of the Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

in a MgB system hydrogen MgB cooling can reduce the TOGW Component Weight Efficiency Specifi c Power 10900 lbs compared to a system (lb) (%) (hp/lb) with cryocooling.

11 84 99.98% 25.3 Generator, The BSCCO system Generator Cooler summarized in Table 8 can operate at 50K. This higher 2189 99 . 28 % 13.7 Generator with Cooler source temperature for the Transmission line cryocooler reduces the weight of the cryocoolers for the motor 200 99.93% 20 Inverter and generator by about half 67 99.57% Inverter Cooler compared to the MgB system.

267 99.50% 15 While hydrogen cooling would Inverter with Cooler have the same advantages with 314 99 . 97 % 13.4 Motor the BSCCO system as the MgB Motor Cooler system, the impact wouldn’t be as significant. For this reason, 516 98 . 95 % 7.8 Motor with Coole r the N3-X with cryocooling uses 3.5 Total - Cryocooled 17123 97.75% the weights and efficiencies defined for BSCCO based 5.4 Total - LH2 Cooled 11078 99.88% motors and generators.

The TeDP systems of the N3-X Table 7 MgB Based Electrical System Weights and Efficiencies.

vehicle, even with the additional BSSCO weight of the electrical Component Weight Efficiency Specifi c Power transmission system, are lighter (lb) (%) (hp/lb) than the total propulsion system weight of the pylon mounted 954 99.93% 31.4 Gener ator, UHB engine on the N3A with Generator Cooler the same assumed technology level. Contributing to the lower 1534 99.55% 19.6 Generator with Cooler weight is the improved specific Transmission line fuel consumption of the TeDP 200 99.93% 20 Inverter relative to the UHB resulting in less fuel burn, which allowed 67 99.57% Inverter Cooler the N3-X to be smaller and 267 99.50% 15 Inverter with Cooler lighter than the N3A aircraft.

This in turn reduced the thrust 298 99.94% 13.4 Motor required of the TeDP engine, Motor Cooler allowing the engine to be made 391 99.48% 10.2 smaller. The end result is that Motor with Cooler the smaller fan area (spread over 4.3 Total - Cryocooled 13938 98.54% 15 small fans rather than 2 large 5.8 Total - LH2 Cooled 10378 99.80 % fans) and smaller core engines reduced the turbomachinery weight 13552 lbs. The Table 8 BSCCO Based Electrical System Weights and Efficiencies.

embedded design of the TeDP saves 10348 lbs in the inlet, nacelle and bypass nozzle with cryocooling. The combination of lower weight and eliminates the pylon entirely. The empty weight and lower fuel weigh means that Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

Figure 11 contains the decomposition Engine Weights Referenc N3A TeDP TeDP of the mission fuel burn reduction for (total for the e Engine ( lb) (Cryo) (LH2) the N3A and the N3-X with vehicle) (lb) (lb) (lb) cryocooling and liquid hydrogen cooling giving the fuel burn reduction Turbomachinery 33622 28887 15335 15335 attributed to each technology applied.

(core and fan) Gearbox/Electric 0 592 13938 11841 Conclusions al The hybrid wing body (HWB) aircraft combined with turboelectric Inlet/Nacelle/ 8829 13377 3029 3029 distributed propulsion (TeDP) system Nozzle /Pylon is able to reduce the mission fuel burn by 70%-72% from that of a B777- Propulsion 42451 42856 32302 30205 200LR-like vehicle (block fuel burn System Total of 279800 lbs), without Weight compromising payload, range or cruise speed. This is accomplished by Table 9 Engine weight comparison using an electrical drive system that decouples the power producing parts Reference N3A N3 - X N3 - X of the system from the thrust Aircraft (lb) (C ryo) (LH2) producing parts of the system with (lb) (lb) (lb) only a relatively lightweight and Empty Wt 340800 285800 26780 267400 flexible electric transmission lines Payload 118100 118100 118100 118100 connecting them. This freedom to Wt configure and locate those two major Total Fuel 309800 147200 93400 88000 portions of the propulsion system Wt allowed each to be optimized for its Block Fuel 279800 1337 00 83500 78500 task.

Wt Fifteen propulsors were located in a Reduction 52.2% 70.2 % 71.9 % continuous nacelle with 2- D “mail - in Block slot” inlets and nozzles that covered Fuel the entire 60 foot span of the center TOGW 768700 551000 479300 473500 body near the upper surface trailing edge of the N3-X aircraft. This Table 10 Aircraft Weight Comparisons and Percent Block Fuel maximized the amount of boundary Burn Reduction layer that was ingested by the system.

Despite the wide span of the total large differences in the configuration of the two array, the aspect ratio of each individual engines make it difficult to single out aspects propulsor inlet is only 2 to 1. The aspect ratio of that lead to this difference in weight. Aspects, each nozzle is a similar 2.7 to1. This allows such as the 2-D nozzle of the TeDP propulsors short, minimal offset, low loss inlets and nozzles that allow a variable fan nozzle area to be with only a fan and motor between. The short accomplished with a simple hinged flat nozzle axial length of the propulsor allows placement of flap, certainly contribute to the weight the inlet at the 85% chord location while still difference. More detailed analysis will be keeping the nozzle plane well forward of the needed to understand the differences.

trailing edge to retain the fuselage noise The end result of this analysis is that the N3-X shielding benefit of the HWB configuration. Aft with a TeDP system was able to meet the SFW of the 80% chord location the inviscid portion of project goal of 70% mission fuel burn reduction.

the inlet flow has less than freestream velocity Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

Figure 11 Fuel Burn Reduction Brea kdown due to diffusion on the aft portion of the HWB MgB currently the best candidate material.

center body. The ability to keep the inlet aft of Preliminary research results in the areas of the 80% chord location allows the combination cryocoolers, refrigeration, and cryogenic power of BLI and aft diffusion to reduce inlet velocities inverters show that target power to weight goals further than just BLI alone would. This results in may actually be within reach of current a total of 10% reduction in inlet velocity at the technology. Liquid hydrogen cooling presents an 85% chord location at the ADP flight condition. alternative to mechanical refrigeration, and also has other advantages such high heating value Two large turbogenerators that produce the which allows a single pound of hydrogen to power to run the propulsors were located at the displace nearly 3 pounds of jet fuel, reducing the wing tips where they would receive undisturbed mission fuel weight even further. The need to freestream air. The combination of the largest generate the hydrogen may offer synergies with possible turboshaft engine size and a freestream renewable power sources such as solar and wind inlet allowed the highest possible thermal energy. The efficiencies afforded by electrical efficiency for a given set of engine technology power transmission result in very significant fuel assumptions. However all of the thermodynamic burn savings. So significant in fact that the advantage would be retained if the resulting reductions in aircraft size and weight turbogenerator were moved to an inboard lead to TeDP engines that are lighter, with the location, such as the wing root.

weight of their electric system included, than the Many of the technical issues involving conventional pylon mounted turbofan engines, development of flight-weight superconducting including nacelle and pylon, of the same motors, generators and transmission lines appear technology level.

to be addressable and will benefit from very With the ability to meet the fuel burn goal active research on both aerospace and terrestrial established, analysis will continue to determine applications. The key area of investigation is the fuel burn sensitivities to changes in the development of AC tolerant stator designs, with assumed technology levels. This will indentify Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

both the technologies to which fuel burn is most sensitive and the threshold values of those 2035,” NASA Contractor Report CR -2010-216691, technologies required to maintain the 70% fuel burn reduction. A noise analysis will also be McCullers, L. A.: “FLOPS Weight Module performed to determine the aircraft noise with Documentation, Wate.doc,” FL OPS Users Manual, respect to the N+3 noise reduction goal of -71 updated April 2008 db. Lastly, the effect on the lift coefficient of the upper surface suction produced by the propulsor NPSS User Guide Software Release: NPSS_1.6.5 array, and the impact that has on balance field Liebeck, R., “Design of the Blended Wing Body length, will be determined.

Subsonic Transport”, Journal of Aircraft , Vol. 41, No. 1, Jan-Feb. 2004. pp. 10-25 Acknowledgements Thomas,R, Burley, C., Olson, E., “Hybrid Wing The authors would like to thank Mr. Michael Body Aircraft System Noise Assessment With Tong, our “fifth” author , for taking one for the Propulsion Airframe Aeroacoustic Experiments”, team. We would like to thank Dr. Rubén Del AIAA-2010-3913, 2010.

Rosario for his encouragement and advocacy as Tillman, Greg, et all, “Robust Design for we traveled through undiscovered territory while Embedded Engine Systems – BLI Inlet and developing concepts for this new class of Distortion- Tolerant Fan Design”, NASA Contract engines. We would also like to thank Dr.

Number NNC07CB59C, 2010 Richard Wahls, Mr. Gregory Follen, Mr.

Patterso n, J.C., Flechner, S.G., “An Exploratory William Haller and the rest of the SFW Wind-Tunnel Investigation of the Wake Effect of a management team for their assistance and Panel Tip-Mounted Fan-Jet Engine on the Lift- guidance; Mr. George Stefko for his enthusiasm Induced Vortex”, NASA TN D -5729, May, 1970 and advocacy. And we would like to thank Dr.

Kawai, Ron, et al, “Acoustic Prediction Fayette Collier for his leadership when we Methodology and Test Validation for an Efficient started looking at the seemingly impossible Low- Noise Hybrid Wing Body Subsonic Transport”, possibility of electric distribution of main NASA Contract Number NNL07AA54C, 2008 propulsion power in large transport class aircraft. de la Rosa Blanco, E., Hall, C., Crichton, D., “Challenges in the Silent Aircraft Engine Design”, th AIAA-2007-454, 45 AIAA Aerospace Sciences References Meeting in Reno, NV, January 8, 2007 Felder, J, Kim, H. D., Brown, G., Chu, J. “An NASA Research and Technology Program and examination of the Effect of Boundary Layer Project Management Requirements, NASA Ingestion on Turboelectric Distributed Propulsion Procedural Requirements 7120.9. Appendix J.

Systems”, AIAA -2011-0300, 49th AIAA Aerospace Technology Readiness Levels (TRLs), February 05, Sciences meeting in Orlando, FL, January 4, 2011 Friedman, D., “Aerodynamic Prediction Greitzer, Edward M., et al, “N3 Aircraft Concept Methodology and Test Validation for an Efficient Designs and Trade Studies,” NASA Contractor Low- Noise Hybrid Wing Body Subsonic Transport”.

Report CR-2010-216794, Volume 1 and 2, , 2010 NASA Contract NNL07AA54C, 2nd Annual Review, NASA Ames Research Center, January 20, 2010 Bruner, Sam, et al., “NASA N3 Subsonic Fixed Wing Silent Efficient Low-Emissions Commercial Guynn, M. D., et. al., “Engine Concept Study for Transport (SELECT) Vehicle Study,” NASA an Advance Single- Aisle Transport”, NASA TM - Contractor Report CR-2010-216798, 2010 2009-215784 Bradley, M. and Droney, C, “Subsonic Ultra Green Brown, G. V., “ Weights and Efficiencies of Aircraft Research: Phase I Final Report”, NASA Electric Components of a Turboelectric Aircraft Contract Number NNL08AA16B, 2010 Propulsion System” , AIAA-2011-0225, presented at 49th AIAA Aerospace Sciences meeting in Orlando, DAngelo, Martin M., et al., “N3 Small Commercial FL, Jan 4-7, 2011 Efficient and Quiet Transportation for Year 2030- Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.

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Xi, H.X., Gong, W.Z., Zhang, Y., Bi, Y.F., Ding, H.K., Wen, H., Hou, B., Xin, Y., “China’s 33.5 m, 35 kV/2 kA HTS AC Power Cable’s Operation in Power Grid”, Physica C, 445 – 448 (2006) 1054 – 1057 “Lightwe ight, Efficient Power Converters for Advanced Turboelectric Aircraft Propulsion Systems ” , Final Report, NASA 2010 Phase 1 SBIR, NNX10CC71P, MTECH Laboratories, LLC, July 29, Jones, S., NASA GRC Reference model for 90000 lb thrust class direct-drive turbofan, 2011 Snyder, C, et al, “Propulsion Investigation for Zero and Near-Zero CO Emissions Aircraft”, NASA/TM - 2009-215487, May, 2009 Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17 U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for government purposes. The copyright owner reserves all other rights.