Document
Turboelectric and Hybrid Electric Aircraft D rive Key
Performance Parameters
Kirsten P. Duffy University of Toledo , Toledo , Ohio, 4 4 606, USA and Ralph H . Jansen NASA Glenn Research Center , Brook Park, Ohio, 44135 , USA NASA has been investigating electrified aircraft propulsion as a means of furthering its goals of reducing fuel burn, emissions, and noise. However, the electric drive components required introduce weight and efficiency penalties at odds with these goals. The purpose of this paper is to propose electric drive spe ci fic power, electric drive efficiency , and electrical propulsion fraction as the key performance parameters for fully turboelectric , partially turboelectric , and parallel hybrid electric aircraft power system s . The impact s of these parameters on overall aircraft performance are investigate d . R ange equations for each aircraft type are described . The benefits and costs that may result from the electrified propulsion system s are enumerated. A breakeven analysis is conducted to find the minimum allowable electric drive specific power and efficiency , for a given electrical propulsion fraction and battery specific energy , that can preserve the range, payload weight , input energy, and ratio of ope rating empty weight to initial weight of the conventional aircraft .
I. Nomenclature D = drag 𝑔 = gravitational constant L = lift 𝑃 = battery output power batt 𝑃 = electrical drive system output p ower elec 𝑃 = fuel output power fuel 𝑃 = propulsive output power prop 𝑃 = turbine engine output power turb 𝑅 = range of aircraft 𝑆𝑒 = battery specific energy batt 𝑆𝑒 = fuel specific energy fuel 𝑆𝑝 = electric drive specific power elec 𝑇 = cruise thrust 𝑣 = c ruise velocity cruise 𝑊 = initial cruise weight of aircraft i 𝑊 = final weight of aircraft f 𝑊 = electric drive weight elec 𝑊 = aircraft fuel weight fuel 𝑊 = payload weight pay 𝑊 = empty weight of aircraft (operating empty weight) OEW 𝛼 = ratio of takeoff to cruise power 𝜂 = efficiency of electric drive system elec Senior Rese arch Associate, Mechanical, Industrial, and Manufacturing Engineering, and AIAA Member.
Electrical Engineer , Aeronautics Mission Office, and AIAA Member.
American Institute of Aeronautics and Astronautics 𝜂 = overall efficiency of the aircraft o 𝜂 = propulsive efficiency of aircraft prop 𝜂 = thermal efficiency of turbine engine therm = electrical propulsion fraction Subscripts: AC = conventional turbofan AirCraft HE = parallel Hybrid Electric aircraft PE = Partially turboElectric a ir c raft TE = fully T urboElectric a ir c raft II. Introduction h ere is substantial interest in the investigation of improvements to aircraft efficiency through the introduction of electrical components into the aircraft propulsion system. In the case of turboelectric and hybrid electric aircraft ,
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the electrical systems can provide unmatched flexibility in coupling the power generation turbine s to the fan propulsors. This flexibility can allow greater propulsion a irframe integration and can result in reduced noise , emissions, and fuel burn . However, the greatly expanded electrical system introduces weight and efficiency burden s at odds with these benefit s . A potentially promising intermediate step between a conventional turbofan aircraft and a fully turboelectric or electric aircraft is a partially turboelectric or hybrid electric pro pulsion system. Initial studies show that a significant aerodynamic benefit can be achieved while only requiring a fraction of the propulsive power to be managed electrically. However, it is difficult to arrive at authoritative conclusions since the aircraft configurations themselves and many of the major electrical system components have yet to be built or verified. A breakeven a nalysis is presented here to elucidate the electrical power system performance requirements necessary to achieve electrified aircraft propulsion , specifically fully turboelectric, partially turboelectri c , and parallel hybrid electric . This first - order anal ysis provides a framework for comparing electric drive system performance factors, such as the electrical efficiency, in the context of aircraft propulsion systems. The value of this analysis is both to guide electrical system component research as well as to provide aircraft configuration researchers with reasonable component expectations.
S imilar parametric analys e s w ere presented previously for a fully turboelectric propulsion system [1] and a partially turboelectric system [2]. T h e current study investigates a broader array of aircraft types, including the fully and partially turboelectric aircraft already addressed, as well as parallel hybrid electric aircraft . In the cases of partially turboelectric and hybrid electric, the fraction of thrust p ower will be var ied between the turbofan engines and electric distr i b u tion to additional propulsors. A key difference between this study and the prior studies is in the breakeven analysis assumptions . Here the input power and ratio of operating empty weig ht to aircraft initial weight are held constant among the aircraft types, in addition to equating the range and payload weight. The other studies held either the initial aircraft weight or the fuel weight to be the same, as well as the operating empty wei ght.
III. Benefits and Costs of Electrified Aircraft Propulsion A. Benefits of Electrified Aircraft Propulsion The turboelectric aircraft propulsion - derived system benefits have been described in previous papers by Jansen et al. [1 - 2], and the main points are now summarized. Higher propulsive efficiency due to increased bypass ratio (BPR), higher propulsive efficiency due to boundary layer ingestion ( BLI ) , and lift - to - drag ratio (L/D) improvements are facilitated by electrified propulsion.
Introduction of an electric drive system between the turbine and fan allows decoupling of their speeds and inlet/outlet areas. With this approach, high BPR can be achieved since any number and size of fans can be driven from a single turbine. Increasing BP R results in improved propulsive efficiency. Also, the speed ratio between the turbine and the fan can be arbitrarily set and varied during operation, thereby removing the physical constraint levied by either direct shaft or geared coupling . As a result, t he fan pressure ratio and the turbine/compressor ratios can be optimized independently. The propulsive efficiency benefits due to higher BPR could be as high as 4 - 8% [ 3, 4 ] .
BLI increases propulsive efficiency by ingesting lower velocity flow near the air frame into the propulsors, reenergizing the wake , and thereby reducing drag. BLI can be implemented on both conventional tube - and - wing as well as hybrid wing body ( HWB ) aircraft. The propulsor is mounted such that the slow moving flow near the aircraft American Institute of Aeronautics and Astronautics is ingested, reenergized, and exhausted where the aircraft wake would have been. The BLI benefits to propulsive efficiency are expected to be 3 - 8 % [ 4 , 5 ] . Combining BPR and BLI propulsive efficiencies listed here yields improvements of 7 - 17%.
Distributed pr opulsion is expected to improve both lift and L/D ratio through wing flow circulation control. The propulsors can be distributed above, below, or e mbedded in the traditional tube and wing configuration. Likewise, HWB configurations can employ fans distribu ted across the upper surface or e mbedded. Improvements in L/D ratio may result in smaller wing area, and reduced drag and weight. The benefits of lift augmentation can be taken in reduced wing area for a given load capacity or shorter takeoff distances. Re duction in wing area reduces wing weight, lowers drag, and thereby imparts fuel savings. Alternatively, the improved lift could be focused on increased climb rate and reduced takeoff distance in order to decrease the noise footprint around the airfield. T he L/ D ratio could be improved by 8% [ 6 ] - 16% [ 5 ].
B. Costs of Electrified Aircraft Propulsion Introducing an electric drive system , with or without batteries , into the aircraft propulsion system will incur penalties in the form of added weight and reduced eff iciency. Here, the electric drive system includes the electric machines, the power management and distribution system, and the thermal system r elated to heat removal in the two prior systems. Specifically, the electric drive system c ould include generator s, rectifiers, distribution wiring, fault protection, inverters, motors, and the thermal control for those components.
NASA is investigating high performance motors and batteries that could make electrified aircraft propulsion viable.
With regard to the electric drive components, NASA is looking to improve both the efficiency and specific power of generators, motors, inverters, and rectifiers. A NASA research announcement has a goal of developing technologies and demonstrating a MW - cla ss motor with efficiency greater than 96% and power density of greater than 13 kW/kg. This is just one component of the electric drive system. The partially turboelectric STARC - ABL (Single - aisle Turboelectric AiRCraft with Aft Boundary Layer propulsor) a ircraft concept assumes those values for the motors and generators, as well as rectifiers and inverters with 19 kW/kg and 99% efficiency. Stacking up all the components for this aircraft, incl uding cables, circuit protection, and thermal management, yield s an ele ctric drive efficiency of 89.1% [ 7 ] .
With regard to batteries, current state of the art Lithium - ion batteries have a specific energy on the cell level of up to 200 Wh/kg . P rojected values in 15 and 30 years are 650 and 750 Wh/kg respectively for L ithium - sulfur , and 950 and 1400 Wh/kg respectively for Lithium - air [ 8 ]. These values have to be de - rated based on depth of discharge, battery structure, and battery management. For comparison, the specific energy of aviation fuel is approximately 12,000 W h/kg.
Clearly the benefits of improved propulsive efficiency from high BPR and BLI, as well as increased L/D must be greater than the costs of electrified aircraft propulsion, and the balance of these benefits and constraints are presented here.
C. Aircraft C oncepts with Electrified Aircraft Propulsion NASA has been investigating several different electrified propulsion systems for aircraft, including fully turboelectric, partially turboelectric, and parallel hybrid electric systems .
The N3 - X concept shown i n Fig. 1 is a 300 - passenger, hybrid wing body aircraft with a fully turboelectric propulsion system , and a d e sign range of 7500 nm i . Turbine engines are located at the wing tips, powering generators.
Electric power is then transmitted through cables to a series of m otor - driven fans located near the trailing edge of the aircraft. This configuration allows for a higher lift - to - drag ratio due to the hybrid wing body, as well as higher propulsive efficiency due to the increase in fan bypass ratio and boundary layer ingestion. This concept, described by Felder, Brown, Kim, and Chu [ 5 ] was conceived as a future generation aircraft to meet NASA’s goal of 70% fuel burn reduction. O ut o f th e 70% overall improvements , 18 - 20% of fuel burn reduction was attributed to the turboelectric propulsion system architecture.
Figure 2 shows the partially turboelectric concept STARC - ABL, which is a 154 - passenger aircraft with a design range of 3500 nmi. This commercial transport concept was developed for notional entry into ser vice in 2035 , and compared to a similar technology conventional configuration by Welstead and Felder [ 9 ]. The propulsion system consists of two underwing turbofans with generators extracting power from the fan shaft and transmitting it to a rear fuselage, axisymmetric, boundary layer ingesting fan. T he power to the tailcone fan is constant and contributes approximately 20% of the thrust at takeoff and about 45% of the thrust at cruise. Analysis in Ref. 9 indicate s that the partial ly turboelectric concept has an economic mission fuel burn reduction of 7%, and a design mission fuel burn reduction of 12% compared to the conventional configuration. It should be noted that subsequent studies have predicted fuel burn reductions that are in the range of 3 - 4%, bu t were not available for referencing at the time of this publication.
American Institute of Aeronautics and Astronautics Figure 3 shows the PEGASUS concept (Parallel Electric - Gas Architecture with Synergistic Utilization Scheme), which is a 48 - passenger parallel hybrid electric aircraft . This concept is described by Antcliff and Capristan [ 10 ]. A detailed analysis of an intermediate parallel hybrid electric concept was performed by Antcliff et al. [ 11 ], which was based on the ATR - 42 - 500 conventional fuel - based aircraft with a range of 6 00 nmi. The analysis included various levels of battery specific energy, which is a critical parameter as battery weight has been shown to be a significant penalty for these types of aircraft. They found that a specific energy of 750 W h/kg was required t o breakeven on total energy, even as the aircraft weight increased over the baseline value.
The N3 - X, STARC - ABL, and PEGASUS concepts will be used as case studies for the breakeven analysis in this study.
Fig . 1 N3 - X concept. Fig. 2 STARC - ABL concept.
Fig. 3 PEGA S US c oncept.
IV. Breakeven Analysis A. Key Performance Parameters and Key Assumptions In order to conduct the breakeven analysis we first define the key performance parameters (KPPs), the key assumptions , and the electrical power system boundary. Then we will formulate range equations for each aircraft type.
Finally, we find the breakeven relationship by implicitly solving for the electric drive specific power and efficiency while holding constant the ratio of oper ating empty weight to initial weight, payload weight, input energy (from fuel and/or batteries), and aircraft flight range. The resulting parametric curves can be used as the top - level requirements for the electrical power system and bounding guidelines fo r further aircraft exploration.
Specifically, the key performance parameters (KPPs) are : Electric drive system efficiency, .
elec Electric drive system specific power, Sp .
elec Electric propulsion fraction for partially turboelectric and parallel hybrid el ectric aircraft, The breakeven assumptions in this analysis used to determine the values of the KPPs include : The ranges of the conventional and electrified aircraft are equal.
The input energy (fuel and/or battery energy) of the conventional and electrified aircraft are equal.
The payload weights of all the aircraft are equal.
American Institute of Aeronautics and Astronautics The ratio of OEW to in itial aircraft weight are equal, where OEW does not include the weights of the electric drive and batteries.
B. Electrified Propulsion System Definitions Each electrified propulsion system will now be described, along with the boundaries of the electric drive system for each case. Figures 4 - 7 show simplified diagrams of the conventional (fuel - based) turbofan, fully turboele ctric, partially turboelectric, and parallel hybrid electric aircraft propulsion systems , respectively . The conventional turbofan system is considered the baseline aircraft system for comparison. The building blocks of the systems are the energy source (fuel and/or battery), the turbine engine, the propulsor, and the electric drive for the electrified propulsion cases. We denote the conventional turbofan aircraft, fully turboelectric, partially turboelectric, and parallel hybrid electric paramete rs with the subscripts AC, TE, and PE, and HE, respectively. Power is denoted by the letter P , efficiency by , specific energy by Se , and specific power by Sp .
Fig. 4 Conventional, fuel - based aircraft propulsion system (AC) .
Fig. 5 Fully turboel ectric aircraft propulsion system (TE) .
Fig. 6 Partially turboelectric aircraft propulsion system (PE) .
American Institute of Aeronautics and Astronautics Fig. 7 Parallel hybrid electric propulsion system (HE) .
The turbine, propulsor, and electric drive have associated thermal ( ) , propulsive ( ) , and electrical therm prop efficiencies ( ) . The fuel power ( P ) , battery power ( P ), turbine engine power ( P ) , electrical power ( P ) , elec fuel batt turb elec and propulsive power ( P ) are defined as output power of the fuel, battery, turbine engine, electric drive, and prop propulsors , respectively. The variables in each of Figs . 4 - 7 illustrate the association between the propulsive subsystems, powers , and efficiencies for each propulsion system. In the partially turbo electric and parallel hybrid electric case s , we must introduce the electrical propulsion fraction, ξ , which we define as the fraction of total aircraft thrust at cruise produced by electrically driven propulsors. When the electrical propulsion fraction is equal to one , all the thrust during cr uise is provided by electrically driven propulsors. The fully turboelectric system is one in which all the thrust throughout the mission, including takeoff and cruise, is provided by electrically driven propulsors.
Therefore the electric drive system will need to be sized accordingly.
The electric drive s pecific p ower ( 𝑆𝑝 ), e fficiency ( 𝜂 ) , and the e lectrical p ropulsion f raction ( ξ ) are proposed elec elec as the three KPPs of the electric drive system in the electrified aircraft. Specific p ower 𝑆𝑝 is the ratio of the rated elec electric drive output power to its mass . Efficiency 𝜂 is the ratio of the output power to the input power of the electric elec drive system. Electrical p ropu l sion f raction ξ is the fraction of total aircraft thrust at cruise produ ced by electrically driven propulsors. These three KPPs will be used to describe electrical power system performance and establish levels of performance necessary.
The boundary of the electric drive system is d efined to lend meaning to the KPP s. For this paper , the boundary will include generators, rectifier s, distribution wiring, fault protection, inverter s , motor s, and the thermal control for those components. The parallel hybrid electric system do es not require generators. Some variants of the electric al drive system may use a subset of these components or alternative layouts . The specific power and electrical efficiency analyzed in this study includes all of the components inside the boundary. Notably, the turbine engine and the propulsors are outside of the electric drive boundary.
A simplified assessment of the relationship between the electric drive system KPPs and the aircraft range and input energy is proposed for top - level aircraft performance comparisons. The range equations are discussed first, then the input energy, and finally the component weights. The breakeven equations are derived for fully turboelectric, partially turboelectric, and parallel hybrid electric aircraft.
C. Breakeven on Range The basis of the analysis is an ex pansion of the tr aditional terms in the Breguet r ange e quation for fuel - based aircraft to include the efficiency and weight of the electric drive system . The range equation for f or battery - powered aircraft from Hepperle [ 1 2 ] is expanded in a similar way. These equations a pply to situations where overall aerodynamic efficiency, the L/D , and flight velocity are constant over the duration of cruise . Although not true for the entire flight envelope, this description is a reasonable approximation for cruise conditions.
We develop range equations of the typical form representing the conventional aircraft and electrified propulsion aircraft configurations concurrently for comparison . The range equations for fuel - based and battery - based aircraft are, respectively, 𝑆 𝑒 𝐿 𝑊 fuel i 𝑅 = 𝜂 ln ( ) (1) fuel o 𝑔 𝐷 𝑊 f American Institute of Aeronautics and Astronautics and 𝑆𝑒 𝐿 𝑊 batt batt 𝑅 = 𝜂 ( ) (2) batt o 𝑔 𝐷 𝑊 i where Se and Se are the specific energies of the fuel and battery, and is the overall efficiency of the propulsion fuel batt o system .
F or fuel - based aircraft, the final aircraft weight, W , is equal to the initial aircraft weight, W , minus the fuel weight f i W . Thus t he fuel - based range equation is fuel 𝑆𝑒 𝐿 1 fuel 𝑅 = 𝜂 ln ( ) . ( 3 ) fuel o 𝑔 𝐷 1 − W 𝑊 ⁄ fuel i Note that for small values of W / W , fuel i 1 𝑊 fuel ( 4 ) ln ( ) ~ , ⁄ 1 − W 𝑊 𝑊 fuel i i which shows that Eq s . 1 and 2 have a similar form. Th us th e range is approximately proportional to the ratio of the energy source weight to the aircraft initial weight. Since Se << Se , battery weight for the same range will be batt fuel much larger than fuel weight.
T he overall efficiency of each aircraft type is defined in Eqs. 5 - 8 in Table 1 as functions of propulsive efficiency, prop , thermal efficiency, therm , and electric drive efficiency, elec . Note that the propulsive efficiency prop defined here is actually the product of transfer efficiency and propulsive efficiency.
To see how adding the electric drive system affects overall efficiency, the ratio of electrified aircraft to baseline conventional overall efficiency is plotted in Fi g . 8 as a function of electric propulsion fraction. Here it is assumed that the thermal efficiency is 55% and the electric drive efficiency is 90%. Increasing decreases overall efficiency for the turboelectric cases, since the electric drive system is in series with the turbine engine. Since is larger than elec , the hybrid electric system has increasing overall efficiency compared to the baseline. However, the battery therm weight required for hybrid electric will be a significant penalty in the breakeven analysis.
Table 1 Overall Efficiency Equations Aircraft Type Overall Efficiency Conventional Aircraft (AC) 𝜂 = 𝜂 𝜂 ( 5 ) o 𝐴𝐶 prop 𝐴𝐶 therm 𝐴𝐶 Fully Turboelectric Aircraft ( 6 ) 𝜂 = 𝜂 𝜂 𝜂 o 𝑇𝐸 prop 𝑇𝐸 therm 𝑇𝐸 elec 𝑇𝐸 (TE) 𝜂 𝜂 𝜂 prop 𝑃𝐸 therm 𝑃𝐸 elec 𝑃𝐸 Partially Turboelectric Aircraft 𝜂 = ( 7 ) o 𝑃𝐸 (PE) ( ) 1 − 𝜉 𝜂 + 𝜉 elec 𝑃𝐸 𝜂 𝜂 𝜂 prop 𝐻𝐸 therm 𝐻𝐸 elec 𝐻𝐸 Parallel Hybrid Electric 𝜂 = ( 8 ) o 𝐻𝐸 Aircraft (HE) ( ) 1 − 𝜉 𝜂 + 𝜉 𝜂 elec 𝐻𝐸 therm 𝐻𝐸 American Institute of Aeronautics and Astronautics 2.0 1.5 Fully Turboelectric oAC Partially 1.0 / oE Turboelectric Fully Electric 0.5 Parallel Hybrid Electric 0.0 0.00 0.25 0.50 0.75 1.00 Electric Propulsion Fraction Fig . 8 Ratio of electrified to conventional aircraft overall efficiency.
D. Breakeven on Input Energy The input energy of fuel is simply the product of the specific energy of the fuel and the fuel mass . Similarly, the input energy of the battery is simply the pro duct of the specific energy of the b attery and the battery mass . Thus, t he input energy equations are 𝑆𝑒 fuel 𝐸 = 𝑊 ( 9 ) fuel fuel 𝑔 and 𝑆𝑒 batt 𝐸 = 𝑊 . ( 10 ) batt batt 𝑔 E. Relationship among Aircraft Component Weights The final part of the breakeven analysis relates the specif i c power of the electric propulsion system to the other component weights . We know that the initial aircraft weight is defined as the sum of the OEW, payload weight, fuel weight, electric drive system weight (for electrified ai rcraft), and battery weight (for HE aircraft) : 𝑊 = 𝑊 + 𝑊 + 𝑊 + 𝑊 + 𝑊 .
( 11 ) i OEW payload fuel elec batt From Eq. 1 1 we can see that 𝑊 𝑊 𝑊 𝑊 𝑊 elec OEW fuel batt payload ( 12 ) = 1 − − − − .
𝑊 𝑊 𝑊 𝑊 𝑊 i i i i i noting that the payload weight and the ratio of OEW to initial aircraft weight are constant among the aircraft.
For the TE aircraft, where all the power must pass through the electric drive system, the Sp will be defined based elec on the takeoff power rather than the cruise power . If we denote the ratio of takeoff to cruise power as , then the electric drive system weight ratio is [1] 𝑊 𝛼𝑣 elec 𝑇𝐸 cruise = .
( 13 ) 𝐿 𝑊 i 𝑇𝐸 ⁄ ( 𝜂 ) 𝑆𝑝 𝑔 prop elec 𝐷 𝑇𝐸 Alternatively, it is assumed for the partially turboelectric and parallel hybrid electric cases that the electric propulsion power , which is the product of and the propulsion power, is not required for takeoff , so the electric d rive system weight ratio is defined as [2] American Institute of Aeronautics and Astronautics 𝑊 𝜉𝑣 elec 𝐻𝐸 , 𝑃𝐸 cruise = .
( 1 4 ) 𝐿 𝑊 i 𝐻𝐸 , 𝑃𝐸 ⁄ ( 𝜂 ) 𝑆𝑝 𝑔 prop elec 𝐷 𝐻𝐸 , 𝑃𝐸 V. Breakeven Results A. Fully Turboelectric Aircraft (TE) E quations for the fully turboelectric a ircraft are as follows, for the range, input energy, and component weight equations , respectively : 𝐿 ( 𝜂 𝜂 ) 𝑊 prop therm 𝑊 fuel 𝑇𝐸 𝐷 fuel 𝐴𝐶 𝐴𝐶 ln ( 1 − ) = ln ( 1 − ) (1 5 ) 𝐿 𝑊 𝑊 i 𝑇𝐸 i 𝐴𝐶 ( 𝜂 𝜂 𝜂 ) prop therm elec 𝐷 𝑇𝐸 𝑊 fuel 𝑇𝐸 ( ) 𝑊 𝑊 i 𝐴𝐶 i 𝑇𝐸 = (1 6 ) 𝑊 𝑊 fuel 𝐴𝐶 i 𝑇𝐸 ( ) 𝑊 i 𝐴𝐶 and 𝑊 𝑊 𝑊 𝑊 𝑊 𝑊 elec 𝑇𝐸 fuel 𝑇𝐸 OEW i 𝐴𝐶 fuel 𝐴𝐶 OEW = ( 1 − − ) − ( 1 − − ) . (1 7 ) 𝑊 𝑊 𝑊 𝑊 𝑊 𝑊 i 𝑇𝐸 i 𝑇𝐸 i i 𝑇𝐸 i 𝐴𝐶 i Several observations can be made from Eqs. 1 5 - 1 7 . First, Eq. 1 5 shows that the fuel fraction for the turboelectric aircraft will be reduced if the product of L/D and overall efficiency is i ncreased compared to the baseline aircraft.
Then Eq. 1 6 shows that the aircraft weight will increase compared to the baseline, wh ich is a result of the added electric drive system.
To solve this set of equations for Sp , we first assume a value of (e.g. = 100%). Then Eq. 15 will yield elec elec elec the fuel fraction W / W , given the baseline fuel fraction, and assumed values for L/D and . From Eq. 16 we fuel TE i TE find the ratio of conventional initial aircraft weight to turboelectric initial aircraft weight, which is then substituted into Eq. 17 to give the electric drive system w eight ratio W / W . Finally Eq. 13 is solved for Sp . This is elec TE i TE elec repeated for a range of values of , resulting in a curve of vs. Sp for the turboelectric system. This procedure elec elec elec is used in a similar way for the partially turboelectri c and parallel hybrid electric propulsion systems, using the appropriate equations for those aircraft.
Similar to the study by Jansen et al. [1], the electric drive specific power and efficiency required to breakeven on range and input energy were determin ed, based on expected propulsive improvements. Again, the difference between this analysis and the previous analysis is in the breakeven based on constant input energy and constant ratio of OEW to initial weight versus constant initial weight and OEW in t he previous study.
The turboelectric aircraft studied here is based on the NASA N3 - X hybrid wing body fully turboelectric aircraft.
In Felder et al. [ 5 ], the N3 - X was compared to two different baseline aircraft configurations – a conventional tube - and - w ing aircraft (777 - 200LR) and an intermediate hybrid wing body aircraft with conventional propulsion (N3A).
Table 2 details the parameters used in the analysis. For all the aircraft, it is assumed that the transfer efficiency is 80% (which is multiplied b y the propulsive efficiency given in the paper to give prop ) , and the thermal efficiency therm is assumed to be 55%.
First, we look at the effect of aero and propulsive benefits on the breakeven curves. Here t he baseline parameters and L/D are based on the Boeing 777 aircraft , and the maximum benefits are those for the fully turboelectric N3 - prop X aircraft . We look at three benefit levels between the baseline 777 and N3 - X ; these include combined aero and propulsive benefits of 7 %, 1 8%, and 29 % for minimum, medium, and maximum benefits, respectively. The 29% benefit is representative of the N3 - X versus the 777 baseline with the L/D and improvements shown in Table 2 .
prop American Institute of Aeronautics and Astronautics Figure 9 shows the breakeven curves for the three levels of propulsive benefits . Electric drive systems with performance above each curve should result in lower fuel burn. Clearly, improving L/D and leads to lower prop demands on the electric drive system. Table 2 includes the specific power and efficiency expected of a superconducting electric drive system, 7.1 kW/kg and 98.54%. With these values, only the medium and maximum benefits case would result in lower fuel burn. Relaxing the efficiency to 90%, as for a non - superconducting electric drive system, only the maximum benefits case would result in lower fuel burn.
Table 2 Fully Turboelectric Aircraft Parameters Baseline Baseline Turboelectric Parameter 777 N3A N3 - X 2.0 1.8 v (m/s) 255 255 255 cruise W / W 36% 24% fuel A C i AC W / W 48% 54% 48% / 54% OEW i L/D 19 22 22 69.6% 72.2% 77.1% prop Sp (kW/kg) 7.1 elec 98.54% elec Figure 10 shows the ratio of electric drive weight to initial turboelectric aircraft weight as a function of specific power. Clearly, the better the specific power is, the lighter the electric drive system will be. For the minimum allowable specific power of 3. 1 kW/kg for maximum benefits at 100% efficiency, the electric drive system comprises 9.6 % of the aircraft weight. This number quickly falls with increasing specific power. Finally, Fig . 11 shows the increase in the turboelectric aircraft weight as a function of electric drive specific power. This particular breakeven analysis results in heavier aircraft, but with the sa me fuel burn as the baseline aircraft.
100% 10% 9% 95% elec 8% Maximum Benefits 90% 7% Medium Benefits 85% iTE 6% Minimum Benefits /W 80% 5% elecTE 4% 75% W 3% Maximum Benefits - 29% 70% 2% Medium Benefits - 18% Electric Drive Efficiency 65% 1% Minimum Benefits - 7% 60% 0% 0 5 10 15 20 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) Electric Drive Specific Power Sp (kW/kg) elec elec Fig . 9 Breakeven curves Fig . 10 Electric drive weight ratio for turboelectric propulsion for turboelectric propulsion American Institute of Aeronautics and Astronautics 120% Maximum Benefits 115% Medium Benefits iAC Minimum Benefits /W 110% iTE W 105% 100% 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) elec Fig . 11 Ratio of turboelectric to baseline aircraft weight in breakeven analysis Figures 1 2 and 1 3 show the electric drive breakeven curves for the turboelectric N3 - X versus the baseline 777 and the baseline N3A, respectively. The electric drive efficiency and power indicated by the orange symbols is for a superconducting system, whic h has very high performance . In Fig . 1 3 , we see that the electric drive system used in the N3 - X analysis does not provide fuel burn benefits in this breakeven analysis, even though Ref. 5 showed reduced fuel burn. The discrepancy lies in the breakeven analysis assumptions. Here we are assuming equal input power, which in this case is equal fuel burn. This results in a larger aircraft compared to the baseline N3A. However, the N3 - X aircr aft actually had a 7% lower aircraft weight than the baseline N3A. This illustrates the sensitivity of this breakeven analysis to the key assumptions. However, Fig . 1 3 does clearly indicate the necessity of choosing the high - performance superconducting e lectric drive.
100% 100% 95% 95% elec elec 90% 90% 85% 85% 80% 80% 75% 75% 70% 70% Electric Drive Efficiency Electric Drive Efficiency Break Even N3-X Break Even N3-X 65% 65% 60% 60% 0 5 10 15 20 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) Electric Drive Specific Power Sp (kW/kg) elec elec Fig. 1 2 Break even for N3 - X vs 777 Fig . 1 3 Break even for N3 - X vs N3A American Institute of Aeronautics and Astronautics B. Partially Turboelectric Aircraft (PE) E quations for the partially turboelectric aircraft are as follows, for the range, input energy, and component weight equati ons , respectively : 𝐿 ( 𝜂 𝜂 ) 𝑊 prop therm 𝑊 fuel 𝑃𝐸 𝐷 fuel 𝐴𝐶 𝐴𝐶 ln ( 1 − ) = ln ( 1 − ) (1 8 ) 𝜂 𝜂 𝜂 𝑊 𝑊 𝐿 i 𝑃𝐸 prop therm elec i 𝐴𝐶 ( ) 𝐷 ( ) 1 − 𝜉 𝜂 + 𝜉 elec 𝑃𝐸 𝑊 fuel 𝑃𝐸 ( ) 𝑊 𝑊 i 𝐴𝐶 i 𝑃𝐸 = ( 19 ) 𝑊 𝑊 fuel 𝐴𝐶 i 𝑃𝐸 ( ) 𝑊 i 𝐴𝐶 and 𝑊 𝑊 𝑊 𝑊 𝑊 𝑊 elec 𝑃𝐸 fuel 𝑃𝐸 OEW i 𝐴𝐶 fuel 𝐴𝐶 OEW = ( 1 − − ) − ( 1 − − ) . ( 2 0 ) 𝑊 𝑊 𝑊 𝑊 𝑊 𝑊 i 𝑃𝐸 i 𝑃𝐸 i i 𝑃𝐸 i 𝐴𝐶 i These equations are similar to the fully turboelectric case, except in the definitions of overall efficiency (Eq. 7 vs.
Eq. 6 ) and electric drive system weight (Eq. 1 4 vs. Eq. 1 3 ).
The effect of electric propulsion fraction on required electric drive system performance was examined for the case of the partially turboelectric STARC - ABL aircraft concept . Welstead and Felder [ 9 ] performed a systems study of the STARC - ABL aircraft compared to an N+3 Conventional Configuration (N3CC) baseline conventional fuel - powered turbofan aircraft. Table 3 shows the baseline and partially turboelectric aircraft parameters used in the breakeven analysis. The propulsive efficiency for a CFM56 fan is ass umed to be 80%, which is multiplied by the transfer efficiency of 80% to give 64%. Similarly, the propulsive efficiency of 93.9% for the GE hFan is used for the STARC - ABL analysis, and is multiplied by 80% to give 75.1%.
If we assume that L/D and prop ar e constant with changing electric propulsion fraction, then the breakeven curves are as shown in Fig . 1 4 . The STARC - ABL aircraft has an electric propulsion fraction of 45% at cruise, and if we assume that the aero and propulsive parameters L/D and for the STARC - ABL in Table 3 scale with then the prop breakeven curves are as shown in Fig . 1 5 . This shows the effect of the benefits versus the costs of the electric drive system, and the importance of predicting those benefits in this type of analysis.
Assuming constant prop and L/D, Fig. 1 6 shows the electric drive weight ratio; the weights are lower for partially turboelectric compared to the fully turboelectric , since the electric drive system is sized based on cruise power rather than takeoff power . Figure 1 7 shows the ratio of partially turboelectric aircraft to conventional aircraft initial weights .
Table 3 Partially Turboelectric Aircraft Parameters Partially Baseline Parameter Turboelectric N3CC STARC - ABL 45% v (m/s) 206 206 cruise W / W 17% fuel AC i AC W / W 57% 57% OEW i L/D 21.4 22.3 prop 64% 75.1% Sp elec (kW/kg) 2.0 90% elec American Institute of Aeronautics and Astronautics 100% 100% 75% Electric 50% Electric 95% 95% elec elec 25% Electric 90% 90% 85% 85% 80% 80% 75% Electric 50% Electric 75% 75% Electric Drive Efficiency Electric Drive Efficiency 25% Electric 70% 70% 0 5 10 15 20 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) Electric Drive Specific Power Sp (kW/kg) elec elec Fig . 1 4 Breakeven curves for Fig . 1 5 Breakeven curves for partially turboelectric aircraft with partially turboelectric aircraft with constant aero and propulsive benefits. scaled aero and propulsive benefits.
6% 112% 75% Electric 5% 110% 75% Electric 50% Electric 50% Electric 4% 108% iPE 25% Electric iAC 25% Electric /W /W 3% 106% iPE elecPE W W 2% 104% 1% 102% 0% 100% 0 5 10 15 20 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) Electric Drive Specific Power Sp (kW/kg) elec elec Fig . 1 6 Electric drive weight ratio Fig . 1 7 Ratio of partially turboelectric to for partially turboe lectric aircraft with baseline aircraft weight with constant constant ae ro and propulsive benefits. aero and propulsive benefits.
Figure 1 8 shows the results of the breakeven analysis for the STARC - ABL concept at its design electric propulsion fraction of 45% . Here we see that the electric drive effici ency and specific power used in Ref. 9 does result in an aircraft with lower fuel burn. Unlike the N3 - X example, the STARC - ABL aircraft actually has a 3% higher initial weight than the baseline, whereas the breakeven analysis shows a 7% higher initial wei ght at Sp = 2 kW/kg. In elec general, th e breakeven analysis assumptions are similar to the systems study in Ref. 9 ; therefore the results are similar.
American Institute of Aeronautics and Astronautics 100% elec 95% 90% 85% 80% Electric Drive Efficiency Break Even STARC-ABL 75% 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) elec Figure 1 8 . Breakeven for STARC - ABL vs. N3CC.
C. Parallel Hybrid Electric Aircraft (HE) E quations for the parallel hybrid electric aircraft are as follows, for the fuel range, electrical propulsion fraction , inp ut energy, and component weight equations , respectively : 𝐿 ( 𝜂 𝜂 ) therm prop 𝑊 𝑊 𝐷 fuel 𝐻𝐸 𝐴𝐶 fuel 𝐴𝐶 ( ) (2 1 ) ln ( 1 − ) = 1 − 𝜉 ln ( 1 − ) 𝐿 𝑊 𝑊 i 𝐻𝐸 i 𝐴𝐶 ( 𝜂 𝜂 ) therm prop 𝐷 𝐻𝐸 𝑊 𝜉 𝑆𝑒 𝜂 𝑊 batt 𝐻𝐸 fuel therm 𝐻𝐸 fuel 𝐻𝐸 (2 2 ) = ( ) 𝑊 1 − 𝜉 𝑆𝑒 𝜂 𝑊 i 𝐻𝐸 batt elec 𝐻𝐸 i 𝐻𝐸 𝑊 𝑊 batt 𝐻𝐸 fuel 𝐻𝐸 𝑆𝑒 ( ) + 𝑆𝑒 ( ) batt fuel 𝑊 𝑊 𝑊 i 𝐴𝐶 i 𝐻𝐸 i 𝐻𝐸 = ( 2 3 ) 𝑊 𝑊 fuel 𝐴𝐶 i 𝐻𝐸 𝑆𝑒 ( ) fuel 𝑊 i 𝐴𝐶 and 𝑊 𝑊 𝑊 𝑊 𝑊 𝑊 𝑊 elec 𝐻𝐸 fuel 𝐻𝐸 OEW batt 𝐻𝐸 i 𝐴𝐶 fuel 𝐴𝐶 OEW = ( 1 − − − ) − ( 1 − − ) . ( 2 4 ) 𝑊 𝑊 𝑊 𝑊 𝑊 𝑊 𝑊 i 𝐻𝐸 i 𝐻𝐸 i i 𝐻𝐸 i 𝐻𝐸 i 𝐴𝐶 i The additional equation in this case, Eq. 22, results from the assumption that the battery - powered portion of the thrust is defined by the electrical propulsion fraction, . We can see from Eq. 2 2 that the ratio of battery weight to initial aircraft weight is directly proportional to the ratio of fuel specific energy to battery specific energy. The fuel specific energy i s approximately 12,000 Wh /kg, compared to projected battery specif i c energy of 500, 750, or 1000 Wh /kg. It is easy to see that the battery weight can become quite large, making hybrid electric configurations more difficult to implement than partially turboelectric configurations , despite the better overall efficiency . However, there are some conditions under which the hybrid electric configuration is more successful. To that end, we investigate the effect of range, Se , and electric propulsion fraction on the breakeven curves.
batt A breakev en analysis was performed for the parallel hybrid electric aircraft described by Antcliff et al . [ 10, 11 ] .
This is a short - range aircraft devised for 48 passengers; the shorter range makes it a better choice for hybrid electric.
The baseline conventional aircraft is the ATR 42 - 500, which utilizes two turboprop engines. There is an intermediate American Institute of Aeronautics and Astronautics parallel hybrid electric concept with a range of 600 nmi, and the parameters shown in Table 4. Here the propu ls ive efficiencies are calculated assuming a transfer efficiency of 80% and = 55%. The parallel hybrid electric therm PEGASUS concept has a 400 nmi range, and a fully electric (at cruise) PEGASUS concept has a 200 nmi range.
To start, the effect of aircraft range was examined. The aircraft range is approx imately proportional to the baseline aircraft fuel fraction, W /W . Therefore, examining the effect of W /W in the breakeven analysis is fuel AC i AC fuel AC i AC essentially the same as examining the effect of the range. We looked at two values of baseline fuel fraction, W fuel AC /W i AC = 0.0 5 (shorter range) and W fuel AC /W i AC = 0. 09 1 ( baseline 600 nmi ). Compared to the aircraft in the turboelectric and partially turboelectric studies, this range is quite small. Figure 19 shows the electric drive perform ance required for the two ranges , for Se batt = 750 W h /kg and = 25 % . Clearly, the parallel hybrid electric configuration is a better option for shorter range flights, which was expected. Note that the electrical efficiency required for the shorter range flight is very low. This is a result of the parallel configuration. For constant therm , as long as ( elec · prop ) HE > ( therm · prop ) AC , the overall efficiency will be higher than the ba seline. There are certainly weight penalties, especially for the battery weight, but these can be overcome depending on the aero and propulsive benefits, which are quite high for this case.
Table 4 Parallel Hybrid E lectric Aircraft Parameters Parallel Parameter Baseline Hybrid Electric 25%, 50%, 75% v cruise (m/s) 150 150 W fuel AC / W i AC 9.1% W OEW / W i 64% 64% L/D 11 15 60% 72% prop Se (Wh /kg) 500, 750, 1000 batt Sp (kW/kg) 7.3 elec 90% elec Next the effect of batte ry specific energy was examined for the shorter range W /W = 0.0 5 . Figure 2 0 shows fuel AC i AC the results for = 25% for Se = 500, 750, and 1000 W - hr/kg. As expected, carrying the heavier batteries increases batt the performance required of the electric drive system.
100% 100% 90% 90% 80% 80% 70% 70% 60% 60% 50% 50% 40% 40% 30% 30% 1000 Wh/kg Electric Drive Efficiency Electric Drive Efficiency 20% 20% Design Range 750 Wh/kg Shorter Range 10% 10% 500 Wh/kg 0% 0% 0 5 10 15 20 0 5 10 15 20 Electric Drive Specific Power (kW/kg) Electric Drive Specific Power (kW/kg) Fig . 19 Breakeven curves based on Fig . 20 Breakeven curves based on battery aircraft range , = 25% , Se = 750 Wh/kg specific energy , = 25%, W /W = 0.05 .
batt fuel AC i AC Figure 21 shows the breakeven curves for various values of electric propulsion fraction, for W /W = 0.0 5 and fuel AC i AC Se =750 W h/kg , assuming the aero and propulsive benefits are constant. If we assume that these and L/D batt prop change with , normalizing the benefits to = 50%, then the breakeven curves are as shown in Fig . 22. There is a big American Institute of Aeronautics and Astronautics difference between the two charts, and it clearly illustrates the balance between the aero and propulsive benefits, and the costs of the battery and el ectric drive system.
Returning to the assumption that the aero and propulsive benefits remain constant, Fig s . 23 - 25 show the electric drive weight fraction , the battery weight fraction, and the ratio of hybrid electric aircraft weight to conventional air craft initial weight. Compared to the fully and partially turboelectric aircraft, the hybrid electric aircraft requires significan t added weight.
100% 100% 90% 90% elec elec 80% 80% 70% 70% 60% 60% 50% 50% 40% 40% 30% 30% 75% Electric 75% Electric 20% 20% 50% Electric 50% Electric Electric Drive Efficiency Electric Drive Efficiency 10% 10% 25% Electric 25% Electric 0% 0% 0 5 10 15 20 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) Electric Drive Specific Power Sp (kW/kg) elec elec Fig . 2 1 Breakeven curves based on Fig. 2 2 Breakeven curves based on electric propulsion fraction with aero and electric propulsion fraction with aero and propulsive benefits constant, propulsive benefits scaling with Se = 750 Wh/kg, W /W = 0.05 . Se = 750 Wh/kg, W /W = 0.05 .
batt fuel AC i AC batt fuel AC i AC 10% 9% 75% Electric 8% 50% Electric 7% 25% Electric iHE 6% /W 5% elecHE 4% W 3% 2% 1% 0% 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) elec Fig. 2 3 Electric drive weight ratio with equal benefits Se = 750 Wh/kg, W /W = 0.05 .
batt fuel AC i AC American Institute of Aeronautics and Astronautics 25% 300% 250% 20% 200% iHE 15% iAC /W /W 150% iHE battHE 10% W W 100% 75% Electric 75% Electric 50% Electric 50% Electric 5% 50% 25% Electric 25% Electric 0% 0% 0 5 10 15 20 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) Electric Drive Specific Power Sp (kW/kg) elec elec Fig . 2 4 Battery weight ratio for Fig . 2 5 Ratio of hybrid electric aircraft weight equal benefits , to conventional aircraft weight with equal benefits , Se = 750 Wh/kg, W /W = 0.05 . Se = 750 Wh/kg, W /W = 0.05 .
batt fuel AC i AC batt fuel AC i AC Now we look at the 600 nmi range parallel hybrid electric aircraft described in Table 4 , with a fuel fraction of 0.09 1 . Now we attempt to compare breakeven results with the results in Ref. 11 , which show that the 750 Wh /kg battery approximately breaks even on input power, which is one of our analysis assumptions, making it a good study for comparison. The 500 Wh /kg batt ery increase s total energy, and the 1000 Wh /kg battery decrease s total energy.
Figure 2 6 shows the results for the parallel hybrid electric concept in our breakeven analysis for an electric propulsion fraction of 25% . As expected, the 750 Wh /kg battery breakeven line was relatively close to the electric drive efficiency and specific power used in the systems study, which found nearly equal input power for that configuration. Improving Se to 1000 Wh /kg allows a relaxation in the electric drive perform ance. The breakeven batt analysis did not yield any viable electric drive performance for the 500 Wh /kg battery, as expected.
These results look good; however, increasing the electric propulsion fraction to 50% or higher does not yield feasible electric drive properties in this breakeven analysis , while Ref. 11 did find viable configurations. An inspection of the Ref. 11 results show that the assumption of W / W remaining constant is not true for that study . We made an OEW i assumption that the aircraft would need to be sized up to carry the weight of the added batteries. If the assumption is made that W OEW /( W i - W batt ) remains constant, which is similar to the Antcliff results, then viable electric drive configurations can be found for > 25%.
100% 90% elec 80% 70% 60% 50% 40% 30% 750 Wh/kg 20% 1000 Wh/kg Electric Drive Efficiency 10% Electric Drive 0% 0 5 10 15 20 Electric Drive Specific Power Sp (kW/kg) elec Fig . 2 6 Breakeven for p arallel h ybrid e lectric a ircraft example at = 25% American Institute of Aeronautics and Astronautics VI. C onclusion The electrified aircraft propulsion concepts for commercial transport aircraft include a very wide range of propulsion airframe integration options as well as electric drive train options . Bounding analyses or parametric trade studies can be very useful to help narrow choices for detailed studies as well as guide technology development choices. Specifi c p ower, e fficiency , and e lectric p ropulsion f raction have been proposed as key performance parameters ( KPPs ) for the electric drive system of an electrified aircraft. The boundary of the system is defined between the output shaft of the turbine to the input shaft of the propulsor , and includes the electrical machines, power distribution , any other power components related to propulsion , as well as any thermal systems associated with the power system. E quat ions were developed that compare the benefits and co sts of an electrified air craft propulsion system compared to the baseline conventional aircraft . Some key conclusions include : Fully turboelectric aircraft o T he requirement that the electric drive system must provide power for takeoff results in tougher requirements on specific po wer than for partially turboelectric aircraft.
Partially turboelectric aircraft o Assuming constant aero and propulsive benefits, a higher electric propulsion fraction requires a better performing electric drive system, due to the added weight of the electri c drive system.
o A ssuming propulsive benefits that scale with electric propulsion fraction , a higher electric propulsion fraction relaxes the requiremen ts of the electric drive system, since the higher aero and propulsive be nefits cancel the costs of the el ectric drive system.
Parallel hybrid electric aircraft o Parallel hybrid electric aircraft is better suited to shorter range.
o Improving battery specific energy will make hybrid electric configurations more feasible.
o Assuming constant aero and propulsive benefits, increasing the electric propulsion fraction increases the demands on the electric drive system, to an even larger extent than the partially turboelectric system because of the added battery weight .
o A ssuming p ropulsive benefits that scale with electric propulsion fraction , a higher electric propulsion fraction relaxes the requirements of the electric drive system . Again, the higher aero and propulsive benefits cancel the costs of the electric drive system. Ho wever, the added battery weight makes the benefits less dramatic compared to the partially turboelectric system.
All aircraft o T he breakeven curves are very sensitive to the propulsive benefit assumptions.
o T he breakeven analysis is sensitive to the component weight assumptions. Here it was assumed that the ratio of OEW to initial aircraft weight remains constant. It may be that other component assumptions are better for a given configuration, which could easily be incorporated into the breakeven an alysis.
o In general, a t low specific power, the efficiency of the electric drive system dominates. But increasing specific power above a certain level yields diminishing returns .
Acknowledgments This work is sponsored by the NASA Advanced Air Transport Technologies project , and funded by the U.S.
Government under NASA Contract NNC13TA85T . The methods used in this paper build on an analytical approach developed by Dr. Gerald Brown at NASA Glenn Rese a rch Center for preliminary analysis of weig hts of electr ical drive systems.
References [1] Jansen, R. H., Brown, G . V., Felder, J . L., and Duffy, K. P. , “ Turboelectric Aircraft Drive Key Performance Parameters and Functional Requirements ,” 51st AIAA/SAE/ASEE Joint Propulsion Conference, AIAA Propulsion and Ener gy Forum , Reston, VA, 2015 .
[2] Jansen, R. H., Duffy, K. P., and Brown, G. V., “Partially Turboelectric Aircraft Drive Key Performance rd Parameters,” 53 AIAA/SAE/ASEE Joint Propulsion Conference, AIAA Propulsion and Energy Forum , Atlanta, GA, 2017.
American Institute of Aeronautics and Astronautics [3] Felder, J. L., Kim, H. D., and Brown, G. V., “Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid - Wing - Body Aircraft,” AIAA Paper 2009 - 1132, 2009.
[4] Brown, G. V., “Weights and Efficiencies of Electric Components of a Turboelectric Aircr aft Propulsion System,” AIAA Paper 2011 - 225, 2011.
[5] Felder, J. L., Brown, G. V., Kim, H. D., and Chu, J., “Turboelectric Distributed Propulsion in a Hybrid Wing Body Aircraft,” 20th International Society for Airbreathing Engines , ISABE - 2011 - 1340, Gothen burg, Sweden, 2011.
[6] Wick, A. T., Hooker, J. R., Hardin, C. J., and Zeune, C. H., “Integrated Aerodynamic Benefits of Distributed rd Propulsion,” 53 AIAA Aerospace Science Meeting , AIAA SciTech Forum, Kissimmee, FL, 2015.
[7] Jansen, R. H., Bowman, C., a nd Jankovsky, A., “Sizing Power Components of an Electrically Driven Tail Cone Thruster and Range Extender,” 16th AIAA Aviation Technology, Integration, and Operations Conference , AIAA Aviation Forum, Washington, DC, 2016.
[8] Dever, T. P., Duffy, K. P., Provenza, A. J., Loyselle, P. L, Choi, B. B., Morrison, C. R., and Lowe, A. M., “Assessment of Technologies for Noncryogenic Hybrid Electric Propulsion,” NASA/TP - 2015 - 216588, 2015.
[9] Welstead, J. R., and Felder, J. L., “Conceptual Design of a Single - Aisl e Turboelectric Commercial Transport with Fuselage Boundary Layer Ingestion,” 54th AIAA Aerospace Sciences Meeting , AIAA SciTech Forum, Reston, VA, 2016.
[10] Antcliff, K. R. and Capristan, F. M, “Conceptual Design of the Parallel - Gas Architecture with Syn ergistic Utilization Scheme (PEGASUS) Concept,” 18th AIAA/ISSMO Multidisciplinary Analysis and Optimization Conference , AIAA Aviation Forum, Denver, CO, 2017.
[11] Antcliff, K. R., Guynn, M. D., Marien, T. V, Wells, D. P., Schneider, S. J., and Tong, M. T. , “Mission Analysis and Aircraft Sizing of a Hybrid - Electric Regional Aircraft,” 54th AIAA Aerospace Sciences Meeting , AIAA SciTech Forum, San Diego, CA, 2016.
[12] Hepperle, M., “Electric Flight – Potential and Limitations,” NATO Workshop on Energy Efficient Technologies and Concepts Operation , STO - MP - ACT - 209, October 2012.
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