Electrical propulsion in commercial aircraft may be able to reduce carbon emissions, but only if new technologies attain the specific power,1 weight, and reliability required for a successful commercial fleet. The committee considered six different electric propulsion architectures. As shown in Figure 4.1, one is all-electric, three are hybrid electric, and two are turboelectric:
- All electric
- Hybrid electric
—Series/parallel partial hybrid
These six architectures, which are shown in Figure 4.1, rely on different electric technologies (batteries, motors, generators, etc.) The levels of CO2 reduction associated with the different architectures are a function of the configuration, component performances, and missions. The results of system studies on various architectures are summarized in the following section.
All-electric systems use batteries as the only source of propulsion power on the aircraft.
The hybrid systems use gas turbine engines for propulsion and to charge batteries; the batteries also provide energy for propulsion during one or more phases of flight. As shown in Figure 4.1, with a parallel hybrid system, a battery-powered motor and a turbine engine are both mounted on a shaft that drives a fan, so that either or both can provide propulsion at any given time. With a series hybrid system, only the electric motors are mechanically connected to the fans; the gas turbine is used to drive an electrical generator, the output of which drives the motors and/or charges the batteries. Series hybrid systems are compatible with distributed propulsion concepts, which
1 In this report, “specific power” and “specific energy” refer to power and energy per unit mass, respectively, and “power density” and “energy density” refer to power and energy per unit volume.
use multiple relatively small motors and fans. The series/parallel partial hybrid system has one or more fans that can be driven directly by a gas turbine as well as other fans that are driven exclusively by electrical motors; these motors can be powered by a battery or by a turbine-driven generator.
Full and partial turboelectric configurations do not rely on batteries for propulsion energy during any phase of flight. Rather, they use gas turbines to drive electric generators, which power inverters and eventually individual direct current (DC) motors that drive the individual distributed electric fans. A partial turboelectric system is a variant of the full turboelectric system that uses electric propulsion to provide part of the propulsive power; the rest is provided by a turbofan driven by a gas turbine. As a result, the electrical components for a partial turboelectric system can be developed with smaller advances beyond the state of the art than are required for a full turboelectric system. Because it is relatively easy to transmit power electrically to multiple widely spaced motors, turboelectric and other electric propulsion concepts are well-suited to distributed propulsion for higher bypass ratios, and they provide aircraft design options for maximizing the benefits of boundary layer ingestion (BLI) in the fans.
Turboelectric propulsion research is one of the four high-priority approaches identified in this report for developing advanced propulsion and energy system technologies that could be introduced into service during the next 10 to 30 years to reduce CO2 emissions. As detailed in the section Technology Needs, below, hybrid-electric and all-electric systems are not recommended as a high-priority approach because the committee determined that batteries with the power capacity and specific power required for commercial aircraft at least as large as a regional jet are unlikely to be matured to the point that products satisfying FAA certification requirements can be developed
within the 30-year time frame addressed by this report. The same situation applies to technologies associated with superconducting motors and generators, fuel cells, and cryogenic fuels. All-electric battery-powered airplane configurations will be limited to small aircraft (general aviation and commuter aircraft), which are not a significant source of CO2 emissions compared to larger commercial aircraft. For large commercial aircraft it is likely that fuel cell applications will be limited to secondary systems such as auxiliary power units and starter systems. Considerable improvements in the specific power of batteries and fuel cells will have to be attained before these power sources would be considered for large aircraft. In addition, the net reduction in CO2 emissions from using all-electric or fuel-cell systems is greatly minimized unless the electrical power used to (1) charge the batteries or (2) produce the hydrogen used to power the fuel cells is generated using renewable or low-carbon-emission technologies.
The committee drew upon an extensive list of recent electric and hybrid electric aircraft system studies conducted by industry, government, and academia. The experts who conducted many of the studies briefed the committee, and several committee members directly participated in or monitored some of the studies.
The studies considered by the committee are listed and summarized in Table 4.1. These studies can be categorized in different ways. Most were aircraft conceptual studies where advanced electrical components were assumed to be available in the future, but they varied widely in assumed aircraft size, range, electrical architecture,
TABLE 4.1 System Studies of Aircraft with Electric Propulsion
|Name and Organization||Aircraft||Time Frame||Electric Architecture||Components||Component Performance|
|Boeing SUGARa,b,c||Single-aisle||N + 3||Parallel hybrid||Motor (1.3-5.3 MW)
|N + 4||Parallel hybrid (fuel cells, superconducting, cryogenic fuels, BLI, open fan)||Motor
|Bauhausd,e||Regional and single-aisle||N + 3
N + 4
|NASA N3Xf||Twin-aisle||N + 3
(N + 4)
|Turboelectric (distributed propulsion, BLI, superconducting, cryogenic fuels)||Generator (30 MW), motor (4 MW)||>10 kW/kg @ 98% efficiency (and other combinations)|
|ESAerof||Single-aisle||N + 2
(N + 4)
|Turboelectric (distributed propulsion, superconducting)||Generator
|NASA small aircraftg||General aviation||N + 1||Turboelectric (distributed propulsion with powered lift)
|Generator (<1 MW), motor (<1 MW)
|Name and Organization||Aircraft||Time Frame||Electric Architecture||Components||Component Performance|
|N + 3
N + 3
Auxiliary power unit (fuel cell, cryogenic fuel)
|Airbusi||General aviation||N + 1||All-electric
|Batteries Motor, generator||250-400 Wh/kg
|Single-aisle||N + 3||Series hybrid (dist. prop., BLI)||Batteries||800 Wh/kg|
|Cambridgej||General aviation||N + 1||Parallel hybrid||Batteries||150-750 Wh/kg|
|Single-aisle||N + 3||Parallel hybrid||Batteries||750 Wh/kg|
|NASAf STARC-ABL||Single-aisle||N + 3||Partial turboelectric (BLI)||Generator (1.45 MW), motor (2.6 MW)||13 kW/kg|
|Georgia Techk||Single-aisle||N + 3||Parallel hybrid||Motor (1 MW) Batteries||3-5 kW/kg
a M.K. Bradley, and C.K. Droney, 2011, Subsonic Ultra Green Aircraft Research: Phase I Final Report, s.l, NASA CR-2011-216847.
b M. Bradley and C.K. Droney, 2012, SUGAR Phase II: N+4 Advanced Concept Development, s.l, NASA. NASA/CR-2012-217556.
c M.K. Bradley and C.K. Droney, Subsonic Ultra Green Aircraft Research: Phase II. Volume II: Hybrid Electric Design Exploration, NASA/ CR–2015-218704/Volume II, Boeing Research and Technology, Huntington Beach, Calif.
d C. Pornet, C. Gologan, P.C. Vratny, A. Seitz, O. Schmitz, A.T. Isikveren, and M. Hornung, 2015, Methodology for sizing and performance assessment of hybrid energy aircraft, AIAA Journal of Aircraft 52(1):341-352.
e H. Kuhn, Bauhaus Luftfahrt, Ottobrunn, Germany, Future Technologies and Ecology of Aviation, “Propulsion and Energy Systems to Reduce Commercial Aviation Carbon Emissions: Long-Term Perspectives,” presentation to the committee on September 2, 2015.
f James L. Felder, NASA Glenn Research Center, “Systems Analysis and Integration, Advanced Air Transport Technology Project,” presentation to the committee on December 7, 2015.
g Mark D. Moore, NASA Langley Research Center, “Distributed Electric Propulsion (DEP) Vehicles,” presentation to the committee on September 1, 2015.
h Chuck Lents, United Technologies Research Center, “UTRC Presentation to the Committee on Propulsion and Energy Systems to Reduce Commercial Aviation Carbon Emissions,” presentation to the committee on September 1, 2015.
i Peter Rostek, Project Leader Hybrid Electric Propulsion, Hybrid Electric Propulsion—A European Initiative for Technology Development,” presentation at Electric and Hybrid Aerospace Technology Symposium (E&H ATS), Bremen, November 17, 2015.
j C. Friedrich and P.A. Robertson, 2015, Hybrid-electric propulsion for aircraft, AIAA Journal of Aircraft 52(1):176-189.
k Dimitri Mavris, Georgia Tech Aerospace Systems Design Laboratory, “Briefing to the NRC Low Carbon Aviation Committee,” presentation to the committee, December 10, 2015.
electrical performance, time frame, whether cost and environmental impact were considered, and overall level of detail. Each of these studies contributes to the body of knowledge for electric aircraft and has been used by the committee to identify overall trends and establish electric component performance levels that could enable various types of electric aircraft.
The time frame of each study in Table 4.1 is described using NASA’s N + 1 to N + 4 nomenclature to predict when a given concept will achieve a specified level of technology readiness (TRL 62) and, subsequently, initial operational capability (IOC), which occurs when the first aircraft of a given type is placed in service. The specific time frames are defined as follows:
2 NASA uses technology readiness levels (TRLs) to track the maturity of a new technology under development. TRL 6 is achieved when a system or subsystem model or prototype has been verified in a relevant environment.
- N + 1 = TRL 6 achieved in 2010-2015 and IOC achieved in 2015-2025
- N + 2 = TRL 6 achieved in 2015-2020 and IOC achieved in 2025-2030
- N + 3 = TRL 6 achieved in 2025-2030 and IOC achieved in 2030-2040
- N + 4 = TRL 6 achieved in 2035-2040 and IOC achieved in 2040-2050
These studies often assume optimistic and aggressive technology development that is not currently supported by budgetary commitments. Slower development of key energy storage or electric system components can make the N + 3 time frames slide into N + 4 or beyond. For example, in Table 4.1, N3X technologies are reported as N + 3, but the committee believes it is more likely that these advanced cryogenic electric components belong in the N + 4 time frame. Similarly, one version of the ESAero Turboelectric concept relies on superconducting components and likewise belongs in the N + 4 time frame.
The design and usage optimization of electric propulsion architectures over a range of aircraft designs and series of missions is complex and has been only partially investigated by most of the listed studies. The investigation of highly integrated power systems is a continuation of the trend to more electric power systems, such as the one used in the Boeing 787. The possibilities for improved peak power management, running turbines and generators at optimum efficiency, energy storage, safety through redundancy, and possibly eliminating a separate auxiliary power unit (APU) or emergency ram-air driven turbine are just starting to be investigated.
The studies assumed different levels of electrical energy storage and component performance depending on their time frame of interest and on the assumed rate of technology development. The most important parameters for aircraft systems are specific energy (Wh/kg) for energy storage and specific power (kW/kg) for electrical components.
The studies generally assume that energy storage would be provided by advanced secondary (rechargeable) batteries. Calculations of specific energy take into account assumed depth of discharge, efficiency losses, and the weight of the installed system, including structure to support the cells, battery thermal management, and possibly safety containment measures. For any given battery type, it is likely that such considerations will result in a value for specific energy that is much lower than the specific energy for a battery of that type that has not been adapted for use in aviation. The system studies attempted to account for this, so most of the numbers in Table 4.1 are the effective installed performance levels. However, there are likely to be some inconsistencies between the various studies in the installation assumptions and calculation methods. Energy storage performance is discussed in more detail in the next section. Most studies did not look at energy cost, battery replacement cost, system operating cost, or life-cycle carbon emissions.
Electrical component specific power is not evaluated in a consistent fashion in all the studies in Table 4.1. Components of interest include motors, generators, inverters, controllers, conductors, switches, and thermal management. Most studies include the weight of generators, inverters, and their thermal management components in total power system weight, while separately aggregating motors, controllers, and their thermal management components. There may be different assumptions concerning redundancy and safety in each study. Therefore, deducing needed component specific power (for a motor alone for example) may be difficult. Component performance will be discussed in more detail in the next section.
By reviewing, comparing, and discussing the results from the studies, the committee was able to make a series of general observations and establish representative technology challenges and recommendations for component performance requirements to enable various types of electric and hybrid electric systems.
The potential reduction in CO2 possible with all-electric, hybrid electric, and turboelectric aircraft is generally much less than some of the “headline” numbers claimed in the studies. Many studies compare their projected reductions to different baselines. Often the large numbers quoted (for example, 70 percent fuel burn reduction) include postulated performance improvements arising from improvements in other areas, including aerodynamics, structures, operations, and gas turbines. In addition, the improvements arising from electric propulsion are often in comparison to current aircraft, not to future conventional aircraft of the same time period that could also benefit from many of the improvements from these other areas.
Tables 4.2 and 4.3 were created to relate levels of electrical power component performance requirements to various propulsion architectures and aircraft types. The values in Tables 4.2 and 4.3 can be used to help establish
TABLE 4.2 Electrical System Component Performance Requirements for Parallel Hybrid, All-Electric, and Turboelectric Propulsion Systems
|Aircraft Requirements||Electric Systema||Batteryb|
|Power Capability (MW)||Specific Power (kW/kg)c||Specific Energy (Wh/kg)|
|General aviation and commuter|
|Motor and generator <1||>6.5||n/a|
|Regional and single-aisle|
|Motor 1.5-3; generator 1-11||>6.5||n/a|
|Motor 4; generator 30||>10||n/a|
|APU for large aircraft||Generator 0.5-1||>3||Not studied|
a Includes power electronics.
b Total battery system and usable energy for discharge durations that are relevant to commercial aviation flight times, nominally 1-10 hours. Values shown are for rechargeable batteries; primary (nonrechargeable) batteries are not considered relevant to commercial aviation.
c Conversion factors: 1 kW/kg = 0.61 HP/lb; 1 kg/kW = 2.2 lb/kW = 1.64 lb/HP.
technology targets for technology development planning and to predict the sequential progress of aircraft development as technologies improve. Integrated aircraft propulsion and power systems are enabled relatively early compared to all-electric and turboelectric architectures, and turboelectric architectures are enabled before parallel hybrid and all-electric architectures. Specific technology projections and research targets are discussed in more detail in the next section.
Potential applications (and time frames) for all-electric and parallel-hybrid concepts are based largely on projected advances in energy storage technology. Jet fuel is an excellent way to store energy, with an equivalent specific energy of approximately 13,000 Wh/kg. As shown in Tables 4.2 and 4.3, a regional or single-aisle aircraft is conceivable with batteries having a specific energy of “only” 800 Wh/kg for a parallel hybrid system (or 1,800 Wh/kg for an all-electric system). Even so, these levels far exceed both the current state of the art (200-250 Wh/ kg) and the committee’s projection of how far the state of the art will advance during the next 20 years (400-600 Wh/kg). Of course, smaller general aviation aircraft designed for short-range missions can and have been designed with less-advanced batteries. However, the committee is unaware of any systems studies that show that any electric propulsion system that relies on less-advanced batteries to augment the propulsion system of large commercial aircraft will reduce CO2 emissions more than would a partial turboelectric or conventional propulsion system.
Turboelectric concepts are not dependent on advances in energy storage technologies. However, to be successful they need to take advantage of synergistic benefits achieved through aircraft–propulsion integration, using BLI and distributed propulsion. Turboelectric architectures inherently have lower efficiency than conventional gas turbine propulsion owing to energy conversion and transmission losses, but they can be more readily adapted for boundary layer ingestion and distributed propulsion, which are discussed in detail in Chapter 2. A key benefit of distributed propulsion is the drop in motor size and power required as there are many more motors. This means that smaller (and easier to develop) 1 megawatt (MW) and 2 MW electric motors can be put into service earlier than if fewer larger motors were used. The potential applications (and time frame) for turboelectric concepts will be based largely on projected advances in the specific power of components.
A partial turboelectric architecture (or some other variant of a turboelectric system) is likely to provide the first opportunity for an electric propulsion system to be incorporated in a regional or single-aisle aircraft configuration and is the first application likely to begin having a significant impact on reducing aviation carbon. Areas
TABLE 4.3 Electrical System Components: (A) Current State of the Art of Electric Components for Aircraft Applications, (B) Stated Research Goals for Some Current Research Programs, and (C) the Committee’s 20-Year Projection of the Performance of Electric Components Configured for Aircraft Applications
|Motor and Generator||Power Electronics||Batterya|
|Power Capability (MW)||Specific Power (kW/kg)b||Power Capability (MW)||Specific Power (kW/kg)||Specific Energy (Wh/kg)|
A. Current state of the art
B. Research goalse
NASA 10-year goalsf
NASA 15-year goals
U.S. Air Force 20-year goalsg
Ohio State Univ. 3-year goals
Ohio State Univ. 5-year goals
Airbus 15-year goal
McLaren automotive projectionh
C. Committee’s projection of the state of the art in 20 years (noncryogenic)i
a Total battery system and usable energy for discharge durations that are relevant to commercial aviation flight times, nominally 1-10 hours. Values shown are for rechargeable batteries; primary (nonrechargeable) batteries are not considered relevant to commercial aviation.
b Conversion factors: 1 kW/kg = 0.61 HP/lb; 1 kg/kW = 2.2 lb/kW = 1.64 lb/HP.
c Values shown are for systems currently operational in aircraft.
d Specific power only considers the mass of the motor or generator, not the mass of additional systems required to maintain cryogenic operation. State of the art is for a DC superconductor. Ultimately, a 1.0-1.5 kHz superconductor will be needed.
e Values shown do not necessarily include packaging needed for use on aircraft. For the same level of technology, values of specific power and specific energy that include the weight of required packaging will be lower than those that do not.
f NASA generally matures technology no further than technology readiness level (TRL) 6, meaning that a system or subsystem model or prototype has been verified in a relevant environment. Efforts to achieve the10-year goal are currently targeting TRL 4, meaning that a component and/or breadboard has been validated in a laboratory environment, and the goals for specific power and energy do not include packaging.
g Goals for specific power and energy include packaging needed for use on an aircraft.
h McLaren is a race car producer in the United Kingdom, and these goals are for automotive applications.
i This projection assumes TRL 6 is achieved in 20 years, with advanced systems entering service in 30 years.
for further investigation include determining the sensitivity of partial turboelectric system designs to the specific power of motors and components and trade studies between different conceptual designs. More structurally and aerodynamically efficient configurations, such as those with advanced materials, more laminar flow, higher aspect ratio wings (e.g., truss-braced wing), or hybrid wing–body configurations could reduce energy requirements, thereby making electric propulsion architectures more practical and practical sooner.
Turboelectric propulsion concepts are heavily dependent on advances in aircraft electrical power system technologies. These technologies include generator systems for electrical power generation; power electronics for power conversion, conditioning, and distribution; high power aircraft distribution that includes circuit protection;
motors; and energy storage. A key issue is how to address higher distribution voltages designed for operation at altitude. It is also important to address total aircraft design impacts when assessing different turboelectric propulsion concepts. Limiting analysis to the propulsion system is not a valid approach to comparing different turboelectric propulsion concepts. It is imperative to include subsystems such as thermal management and other aircraft housekeeping electrical power requirements. The Boeing 787 provides the most relevant baseline for a total power system rating of 1 MW with understood size and weight metrics. The Boeing 787 also provides an appreciation of housekeeping (nonpropulsive) power requirements that, when applied to a single-aisle aircraft, will add 500 kW-1 MW of electrical power demand that the turboelectric propulsion system must satisfy.3
Turboelectric propulsion concepts for aircraft are described in the introduction to this chapter. This section includes a discussion of aircraft electrical power technologies and challenges that are critical to enabling turboelectric propulsion concepts. The state of the art and 20-year projections for these electrical power system technologies are summarized in Table 4.3. As indicated there, the goals of many organizations exceed the committee’s projections of how far the state of the art is likely to advance in the next 20 years.
Requirements for electrical system components are beyond the current state of the art, especially for the large aircraft that account for more than 90 percent of global CO2 commercial aircraft emissions. The committee predicts that the state of the art in 20 years will allow requirements for specific power of motors and generators to be met for the lowest end of regional and single-aisle aircraft.
Electric Machines for Aircraft Motors and Generators (Noncryogenic)
Figure 4.2 illustrates the progression of the state of the art for aircraft generators in service and under development with ratings from tens of kW up to MW-class, including wound-field synchronous, permanent magnet, and wire-wound superconducting synchronous machines. Note that power densities for aircraft applications lag the densities of other industrial applications such as automotive and ships, mainly because the stringent operating environment and safety certification requirements unique to aircraft add size and weight to electrical power system components. As indicated, some of the generators depicted were built and tested, but others were not. The generators shown span the range from the current state of the art (2.2 kW/kg) to a future vision with a specific power increased by a factor of ten (22 kW/kg). Many figures of merit are relevant to the selection of aircraft generators. These include dynamic and transient electrical performance, size, weight, efficiency, and reliability. These figures of merit lead to design considerations such as high and variable speeds that result in smaller and lighter machines compatible with variable speed turbine engines. Machine thermal management is also a consideration and must be compatible with aircraft thermal management systems (i.e., excessive demand for cooling fluid/oil to minimize machine size at the expense of an excessive thermal system size will not work).
Hybrid-electric propulsion concepts use generators to make electrical power. These machines deliver torque to perform electric engine start and then become primary sources of electrical secondary (non-propulsive) power once engines are burning fuel. Starter/generator systems will be larger than motor systems due to larger power electronics requirements to provide torque and generate electrical power. Each Boeing 787 has a total of four state-of-the-art generators on the main engines, providing a total of 1 MW of electrical power. The specific power of the generators alone is roughly 2.2 kW/kg. Including associated electronics roughly doubles the weight for a system specific power of 1.1 kW/kg.
Rather than use just one large generator, aircraft use multiple, smaller generators to increase power system reliability for flight-critical systems. For example, the Boeing 787 uses four 250 kW generators, and the F-35 uses two 80 kW generators.
As shown in Table 4.3, the committee projects that the specific power of motors and generators could increase to approximately 9 kW/kg in 20 years, with power levels of approximately 1-3 MW. This could be achieved by increasing machine speed (overcoming today’s limits imposed by mechanical stress), increasing power conversion efficiency (limited by the performance of silicon-based power electronics), and increasing power generation and
distribution voltage (limited by breakdown voltage at altitude). Figure 4.2 depicts many attempts to build MW-class aircraft generators, and the ability to break through this barrier is critical to enabling turboelectric propulsion.
To properly characterize the performance of advanced generators, they must be tested as part of a system that is representative of how they will be used on aircraft. Testing infrastructure available today is limited in its capability to conduct MW-class research and development testing for aircraft applications at both the component and system level.
Power electronics already play a key role for aircraft electrical power systems and that role becomes more critical as those power systems become part of MW-class flight critical turboelectric propulsion systems. Power electronics are used for power conversion (including motor drives) and power distribution (circuit protection). Silicon carbide (SiC) power electronics are enabling for MW-class aircraft power due to their improved efficiency and high voltage performance characteristics compared to today’s silicon-based power electronics. SiC is also a more reliable technology than silicon in commercial aircraft environments. Specific power for silicon-based power electronics systems today is approximately 2.2 kW/kg for aircraft applications, and their use for circuit protection is limited to 25 A at 270 V DC (7kW). Higher powered circuit protection is provided by mechanical breakers and relays up to about 500 A at 270 V DC (135kW) using state-of-the-art equipment. It is envisioned that in 20 years SiC-based power electronics systems for aircraft applications will have a specific power of 9 kW/kg for power conversion and circuit protection using electronic components up to 200 A at ±270 V (essentially 540 V, for a power capacity of 108 kW) or using mechanical breakers up to 1,000 A at ±270 V (540 kW).4 High specific powers will be facilitated by advances in components that make power electronics heavy: switching components,
4 To provide circuit protection, circuit breakers that are rated for a system with these nominal operating parameters will actually need to be able to withstand much higher values of current and voltage to accommodate transients.
materials, switching topologies, passive filter components such as transformers, packaging, and thermal management components.
High Power Distribution
A power distribution system voltage of ±270 V (or 540 V) seems to be the limit for the foreseeable future due to physics-based limits referred to as Paschen curve limits. This voltage is used on the Boeing 787 today, and the U.S. Air Force is investigating the use of ±270 V for future high power aircraft. Many high power turboelectric system concepts include kilovolt-class power distribution systems. Such high voltages would require new types of insulation systems and electrical conductor spacing rules and practices.
Circuit protection and high-power distribution cabling for MW-class aircraft power systems also need to be developed. Every electrical circuit on an airplane must have circuit protection. MW-class circuit breakers may exist for power plants in ground and marine applications, but it should not be assumed that the technology incorporated in these breakers is applicable to aviation unless and until it has been verified that aircraft requirements related to weight, volume, voltage, etc. can be resolved. The committee is not aware of any ongoing circuit protection development for MW-class aircraft power systems. The preceding section, Power Electronics, describes current and future projected capabilities of circuit protection components.
Cryogenic/superconducting power distribution is discussed in a later section, Cryogenic Electric Aircraft Power Systems. Efforts have been made to develop other types of conductors for power transmission (e.g., polymers and nanotubes/graphene), but without any notable successes relevant to aircraft applications in the near- to mid-term.
Power System Efficiency
Electrical power system efficiency claims for many hybrid-electric propulsion concepts are more than 95 percent, but they are for machines only and do not include any type of power conversion. In addition, electrical power efficiencies are generally given for full load conditions. Efficiencies drop off when not at full load due to inherent losses in the system that exist independent of load. Figure 4.3 depicts an example of power system efficiency. Assuming 95 or 99 percent conversion efficiency at each step, as shown, the electrical power drive system (i.e., the components between the gas turbine engine and the propulsor) will have a combined efficiency of about 80 percent (0.95 × 0.95 × 0.99 × 0.95 × 0.95 = 0.8). When this efficiency is combined with turbine engine and propulsor efficiencies, as illustrated in Figure 4.3, total fuel to propulsor efficiency is 35 percent (0.55 × 0.8 × 0.8 = 0.35).5 If benefits assessments of turboelectric systems account for actual electrical system efficiency along with the size and weight of the electrical power system components, then they can be compared with similar analyses of conventional and other propulsion concepts. Some literature has suggested eliminating conversion and control electronics to have the generators directly drive the propulsor motors, but this concept has not yet been demonstrated for applicability to turboelectric propulsion. If successful, this approach would eliminate two power conversion stages in Figure 4.3 (the converter controller and the motor drive).
The ability of aircraft to manage heat will be a limiting factor for the high-power electrical power systems needed for turboelectric propulsion. The thermal management system itself will require electrical power to operate, and that power demand will need to be accounted for along with the demands of other nonpropulsive (secondary)
5 The estimated component efficiencies are intended to be illustrative. Even if ongoing efforts to improve these efficiencies are successful (for example, by improving the efficiency of motor/generators and propulsor motors from 95 percent to 98 percent, total system efficiency would increase from 35 percent to 38 percent. In addition, the estimated efficiency shown here does not account for energy required to operate thermal management systems to remove waste heat resulting from system inefficiencies.
power systems. Approaches for thermal management of MW-class turboelectric aircraft propulsion systems have not been addressed in detail in system trade studies to date.
Batteries have been proposed for powering electric aircraft, either as stand-alone systems or hybridized with other power generation systems. Batteries offer modular building blocks for a wide variety of operational concepts in centralized or distributed power systems. They can respond very quickly to changing power demands and could be used to meet peaking or load-leveling requirements. Batteries provide electrical power with no direct carbon emissions, but the indirect emissions from the power source used to charge the batteries must be taken in account. Electricity generated using fossil fuels has substantial CO2 emissions, but these emissions can be mitigated more effectively on the ground than at altitude. The power industry has access to increasingly efficient and effective means of limiting the release of ground-based CO2 emissions, and if the electricity source is powered by solar, wind, or nuclear power, the indirect emissions are nearly carbon-free.
Many organizations have developed and demonstrated short-range, one- or two-passenger, battery-powered electric aircraft using commercially available batteries. A smaller number of similarly sized hybrid-electric aircraft have also flown. A battery system with specific energy greater than 800 Wh/kg is required to enable parallel hybrid propulsion systems on regional and single-aisle aircraft (see Table 4.2). All-electric regional and single-aisle aircraft would be suitable only for short-range operations, and even then they would require a battery system specific energy of 1,800 Wh/kg. This far exceeds the committee’s 20-year projection for battery specific energy (400-600 Wh/kg).
Lithium-ion batteries currently dominate the market in both consumer electronics and electric vehicles. Batteries can be scaled to meet power and energy requirements for aviation, as lithium-ion battery systems with power capability greater than 10 MW and energy storage capacity greater than 10 MWh have already been demonstrated in stationary energy storage for electric utility applications. The key challenge to battery-powered propulsion systems for aviation is to increase battery specific energy. Current lithium-ion battery cells typically achieve 150-250 Wh/kg, while commercial lithium-sulfur cells currently achieve over 350 Wh/kg.
Advances in cathode materials, anode materials, and electrolytes have the potential to perhaps double the specific energy of lithium-ion batteries. Even higher specific energies may be possible in the longer term through the use of advanced battery concepts, such as lithium-sulfur, aluminum-air, and lithium-air.
The theoretical specific energies, based on the active materials, for lithium-sulfur (2,680 Wh/kg), aluminum-air (8,140 Wh/kg), and lithium-air batteries (11,000 Wh/kg in the charged state; 3,500 Wh/kg in the discharged state) are quite impressive. Practical specific energies are significantly lower, however, due to the added weight of current collectors, electrolytes, separators, battery cases, and terminals. Thus, while batteries with specific energies of 1,500 Wh/kg may be achievable, such high specific energies will require major breakthroughs.6 Furthermore,
6 P.G. Bruce, S.A. Freunberger, L.J. Hardwick, and J. Tarascon, 2012, Li-O2 and Li-S batteries with high energy storage, Nature Materials 11(1):19–29, doi:10.1038/nmat3191.
the requirement to simultaneously achieve long cycle life, low cost, and acceptable safety greatly increases the complexity of the overall challenge.
Major technological innovation in “beyond lithium-ion” battery systems will be required to achieve the range of acceptable specific energies needed for commercial introduction of battery-powered electric and hybrid aircraft propulsion systems before these systems can make a significant contribution to reducing carbon emissions in commercial aviation.
The environmental benefits of all-electric aircraft will be offset by CO2 emissions from the source of electricity used to charge their batteries. In addition, the economic viability of all-electric aircraft will depend, in part, on the cost of new infrastructure, including infrastructure on site at airports to charge aircraft propulsion batteries. In addition, a fleet of large commercial all-electric aircraft would only be possible with new or upgraded power transmission lines to airports and, potentially, new generating capacity.
Fuel cells convert the chemical energy in a fuel into electrical power without any combustion. The exhaust from fuel cells is totally carbon-free if hydrogen is used as the fuel. However, if a hydrocarbon fuel is used, the exhaust still contains CO2 in direct proportion to the amount of fuel consumed, but there are no NOx or particulate emissions.
Two types of fuel cells that have been developed for automobile transportation and stationary power generation applications can be considered for aviation. The proton exchange membrane (PEM) fuel cells operate at 80°C to 120oC and require pure hydrogen as the fuel; if a hydrocarbon fuel is used for them, it will have to be first reformed to produce pure hydrogen without any CO, which easily poisons PEM fuel cells. Solid oxide fuel cells (SOFCs) operate at 750°C to 1000oC and can use a variety of hydrocarbon fuels, including jet fuels.
Fuel cells have been investigated for a variety of aviation applications, including these:
- Auxiliary power units (APUs),
- Low-altitude aircraft propulsive power,
- High-altitude long-endurance aircraft,
- Airport applications,
- Ground support equipment,
- Mobile lighting,
- Mobile generators, and
- Unmanned air vehicles.
Sandia National Laboratories analyzed the use of PEM fuel cells in a commercial aircraft (Boeing 787 Dreamliner) to assess the feasibility of having a fuel cell system on the airplane and the impact on other airplane systems and flight performance. They concluded that a fuel cell system onboard a commercial airplane is technically feasible, that it would perform well electrically, that it would be a flexible power source, and that recovery of heat and/or water would allow the fuel cell system to pay for itself and reduce the consumption of jet fuel.
Pacific Northwest National Laboratory evaluated the use of SOFC APUs in a more electric aircraft; again, the Boeing 787 Dreamliner was chosen for the study. The researchers concluded that the weight of the existing SOFCs would have to be reduced by a factor of 2 or 3 in order to compete on the basis of total fuel consumption during flight, and significant increases in specific power would be required to achieve fuel savings during flight.
A collaboration between Protonex and the Naval Research Laboratory (the Ion Tiger Fuel Cell UAV Demonstration), showed the use of PEM fuel cells in an unmanned air vehicle (UAV) for 26 hr of flight on gaseous hydrogen fuel and 48 hr on liquid hydrogen fuel. The Air Force Research Laboratory demonstrated the use of a SOFC power system utilizing JP-8 logistic fuel in Air Force UAV applications; it concluded that the key challenges are to achieve power system efficiency greater than 30 percent while operating on logistic fuels for extended mission durations (more than 50 hrs) and to integrate the fuel cell system into a package with high specific power (more than 150 W/kg). The U.S. Air Force is developing SOFCs with high specific power (more than 500 W/
kg); current SOFC power systems have a specific power of less than 100 W/kg compared to about 1,000 W/kg for internal combustion engines.
Boeing, with General Electric, has investigated concepts including fuel cell/turbine and fuel cell/battery hybrids (SUGAR N + 4 concepts) to reduce airliner fuel consumption and emissions. Boeing conducted fuel cell demonstrator airplane flight tests in 2007 and 2008, and in 2012 it conducted 737 ecoDemonstrator flight tests with a high-temperature PEM fuel cell with Japan’s Ishikawajima-Harima Heavy Industries. Other organizations (Bauhaus Luftfahrt and DLR, both of Germany) are also exploring similar fuel cell hybrid systems for aviation.
The success of the above efforts notwithstanding, daunting challenges will need to be overcome to use fuel cells as part of an electric propulsion system.
PEM fuel cells are presently being designed and built for automotive and APU applications, generally in 1 to 100 kW sizes; the feasibility of commercial scale-up to the sizes necessary for MW-class commercial aircraft needs to be established. Hydrogen (fuel) storage, either as compressed hydrogen gas or as liquid hydrogen, operation at high altitudes, and transient operating conditions are key challenges and concerns.
Hydrogen storage for hydrogen-fueled fuel cells is a problem in terms of size, weight, thermal management, and airport infrastructure. The weight of the storage systems affects both the specific power and the energy of a fuel cell system, and the size affects both power and energy density. Regenerative fuel cells that make, store, and then consume hydrogen could conceptually be used as an energy storage system for hybrid electric propulsion systems. Regenerative fuel cells, however, are more complex than other fuel cells, and they have low round-trip energy efficiency.
SOFCs are being developed for both large-scale stationary power applications (more than 100 kW) and small-scale (1 to 10 kW) APUs and residential applications. SOFCs work better under consistent, steady power conditions; for aviation applications, transient response times, and on/off thermal cycles need to be improved. A key challenge is to increase specific power from less than 100 W/kg presently to over 500 W/kg for potential aviation applications.
There are no currently certified fuel cell systems on a commercial aircraft. The current technology readiness level for PEM fuel cells and SOFCs is TRL 4/TRL 5.7
Because of the challenges and concerns outlined here, while fuel cells may contribute as a power source for auxiliary power units, the committee does not foresee their contribution as an aircraft propulsion source in the timeline of this study. To be considered as a propulsion source, vast improvements in specific power would have to be achieved.
The development of cryogenic electrical power generation has been pursued for aircraft applications for several decades, and much progress has been made. Even so, there is still work to be done, and it is difficult to assess realizable specific powers of cryogenic generator systems for aircraft applications. Cryogenic machines have been built and tested, but without the power conditioning required for an aircraft generator system; the cryogenic machines as tested are not able to handle any type of transient load. Machine testing typically evaluates only open and short circuit performance to evaluate machine electromagnetics, and this type of testing is not sufficient to demonstrate the suitability of cryogenic generators for high-power aircraft applications.
One of the persistent long-term challenges for cryogenic electric aircraft power systems is the lack of superconducting material suitable for alternating current (AC) applications; AC-tolerant superconductors have the potential to reduce the mass and weight of wires by orders of magnitude compared to copper. However, even direct current (%) distribution systems experience transients that will present problems for non-AC compatible material. Without high-frequency AC superconductors available, only the machine field winding, which uses a DC current, is superconducting. To get machines to sizes appropriate for aircraft installation, a 1000-1500 Hz compatible superconductor is required for a full cryogenic generator. Another challenge with superconductors is the need
7 NASA TRL 4 and 5 mean that a component and/or breadboard has been validated in a laboratory environment (for TRL 4) or in a relevant environment (for TRL 5).
to perform voltage regulation; conventional approaches to voltage regulation will not work with a conductor that has zero electrical resistance.
Cryogenic system operating temperature is an important consideration. The more powerful the magnetic field in which a superconducting material operates, the lower the required cryogenic temperature will be. For example, superconducting materials being developed today for operation at 77 kelvin (K) do not function as superconductors in a 10 tesla magnetic field until the material is cooled down to 20 K. Cryocoolers able to cool to 20 K are four times the size and weight of 77 K cryocoolers. Even for an aircraft powered by liquefied natural gas (LNG), given that LNG is stored at 112 K, cryocoolers would be required for cooling from 112 K to 20 K. Any cryocoolers used on aircraft will have to be very robust, with redundant capability on board, to meet safety requirements in the event of a cryogenic component failure.
The minimum capacity of cryocoolers is determined by the thermal load on the cryogenic systems the cooler is designed to support. This may be just a few watts, but cryocoolers sized to meet this minimum requirement will take days to cool an aircraft system from ambient to operating cryogenic temperature, during which time the aircraft would be out of service. Faster cooling methods can be employed, but only if the system is designed to withstand the resulting thermal shock.
In summary, cryogenic technologies have the ability to greatly reduce the specific power of electrical systems for a wide variety of applications, including aviation. However, there are substantial barriers to the implementation of these technologies in the challenging operational environment of a commercial aircraft, and it is not envisioned that technology for cryogenic power generation or power distribution will be ready for incorporation in an aircraft propulsion system within the 30-year time frame addressed by this report.
As discussed in Chapter 1, reductions in carbon emissions from general aviation and commuter airline operations will not make a significant difference in total carbon emissions from global commercial aviation. However, these aircraft types can serve as technology development platforms and will help define an application pathway to regional and larger aircraft. So even though all aircraft categories are described in this section, the focus for significant carbon reductions in the future will have to be on single- and twin-aisle aircraft, but with some consideration also given to regional aircraft.
As shown in Table 4.2, the electrical system requirements for aircraft of a given class will vary according to the type of propulsion system (e.g., parallel hybrid, all-electric, and turboelectric). Of these three, turboelectric systems permit the use of the smallest electrical machines, which reduces technology development risk and schedule. A series/parallel partial hybrid system could also be developed with relatively small electrical machines and batteries, but commercial aircraft studies to date have not shown that such a system is preferable to either a conventional propulsion system or a partial turboelectric system.
All-electric technology aircraft have been operating as technology demonstrators and are starting to enter production. Some examples of these aircraft, which represent the state of the art as of 2015, are shown in Figure 4.4. They use relatively small electric motor systems, on the order of 60-80 kW.
In April 2015, Siemens announced the development of a direct-drive (2,500 rpm), 260 kW aircraft electric motor weighing a little over 100 lb. The motor specific power is on the order of 5 kW/kg, and it is capable of powering aircraft with a maximum takeoff gross weight of 4,000 lb. The potential availability of such an engine suggests that twin-engine commuter aircraft could be powered by electric motors using current technology. Of course, the question remains regarding the weight and performance of the power source, be it batteries or some hybrid system, as these will determine the potential range of such aircraft and hence their economic viability.
Notwithstanding the potential availability of the Siemens electric motor, NASA has been studying an advanced nine-passenger aircraft concept with distributed electric propulsion that could replace aircraft like the nine-passenger Cessna 402. Cape Air operates 84 Cessna 402s in the Northeast, Midwest, Montana, and the Caribbean. Figure 4.5 shows the Cessna 402 alongside an advanced distributed propulsion concept. Arguably, this eight-motor concept could be powered by 60 kW motors, which are already used in the emerging general aviation industry for two-place trainer aircraft.
In 2010, NASA studies were completed on a range N + 3 aircraft with capacities ranging from 20 to 180 passengers. Study teams were led by General Electric, the Massachusetts Institute of Technology, Northrop Grumman, and the Boeing Company. Probably the most thoroughly investigated concept since that initial study is the Boeing SUGAR (Subsonic Ultra Green Aircraft Research) family of aircraft with various electric propulsion architectures. The SUGAR Volt concept features a twin-engine aircraft (see Figure 4.6) and relies upon projected advances in battery technology to enable a parallel hybrid electric propulsion system. The aircraft’s concept engines were provided by General Electric and used a large electric motor attached to the low pressure shaft of a gas turbine. This allowed the turbofan to run conventionally, burning aviation fuel, or it could use the electric motor to augment the power supplied to the low pressure shaft (and hence the fan). The motor could also provide exclusive power
to the fan when the core of the gas turbine is shut down during portions of the cruise mission, thereby significantly reducing emissions. Figure 4.6 shows the primary SUGAR Volt airframe architecture. Another element of the SUGAR project, the SUGAR Freeze, considered the viability of using fuel cells, cryogenic technologies, and boundary layer ingestion (using a propulsor in the aft fuselage).
Empirical Systems Aerospace has developed the ECO-150 concept aircraft in two variants, one that would
use cooled, superconducting technology, and another that would use nonsuperconducting electrical systems. Both concepts would use a turboelectric system with distributed propulsion.
Figure 4.7 shows the Distributed Open Rotor Aircraft electric propulsion concept that has been studied by Rolls-Royce in the United Kingdom. This concept uses multiple propulsors distributed across a significant portion of the wing. These propulsors could be either electrically powered by turbogenerators mounted on the wings or shaft driven. The study evaluated an electrical power distribution system based on electrical technologies available at TRL 4 today. Regardless of the power distribution mechanism used for this concept, the slipstream of the propulsors significantly increases the wing lift coefficient at takeoff, allowing the wing area to be reduced and enabling the wing to be designed with a very high aspect ratio, optimized for highly efficient cruise performance. A key feature of this concept is its use of distributed propulsion (described in Chapter 2). For an electric propulsion system, the use of many fans lowers the power required of each motor. This potentially makes the aircraft practical sooner, as smaller motors can be used instead of waiting for the development of larger motors.
Partial turboelectric systems are also being studied. The NASA STARC-ABL (single-aisle turboelectric aircraft–aft boundary layer) is an example of a partial turboelectric system (see Figure 4.8).
Airbus and Rolls-Royce are studying the e-Thrust concept (see Figure 4.9), which relies on cryogenically cooled superconducting technology that is unlikely to be ready for operational use within the 30-year time frame addressed by this report.
As noted above, no electric propulsion concepts will mature to the point that they can meet the needs of twin-aisle aircraft within the 30-year time frame addressed by this report.
Finding. Rationale for Turboelectric Propulsion Research. Turboelectric propulsion systems are likely the only approach for developing electric propulsion systems for a single-aisle passenger aircraft that is feasible in the time frame considered by this study. System studies indicate that turboelectric propulsion systems, in concert with distributed propulsion and boundary layer ingestion, have the potential to ultimately reduce fuel burn up to 20 percent or more compared to the current state of the art for large commercial aircraft.
Batteries could be a propulsion power source for a small all-electric aircraft (not a subject of this study) within the timeline of this study. To be considered a propulsion source, significant improvements in specific power will have to be achieved and advanced batteries will need to meet economic, safety, and reliability requirements. In particular, the benefits of all-electric aircraft could be offset by CO2 emissions from the source of electricity and the cost of new infrastructure.
While fuel cells may contribute as a power source for auxiliary power units, the committee does not foresee their contribution as a source of aircraft propulsion in the timeline of this study. To be considered as a propulsion source, vast improvements in specific power would have to be achieved.
The state of the art of electrical technologies for motors, generators, power distribution, and power electronics (for example, inverters, converters, and circuit protection) will need to advance to enable turboelectric propulsion concepts for large commercial aircraft.
Electrical machines (motors, generators, inverters) need to be developed to attain specific power, weight, and reliability for commercial aircraft application. Specific power will have to be improved by a factor of between 5 and 10 from the current state of the art. Machine thermal management targets will have to be compatible with aircraft thermal management systems. Electrical machine efficiencies will have to be improved from 95 percent to 97 or 98 percent. It remains to be seen if the voltage of the aircraft power distribution system should be increased as part of the effort to reduce weight and increase specific power.
Circuit protection and high-power distribution cabling for MW-class aircraft power systems will need to be developed.
Advanced materials will likely play a key role in advancing the state of the art of electrical machines and power electronics. Motor efficiencies and power density, for example, can be improved through the use of better conductors, insulators, magnets, bearings, thermally conductive materials, polymer composites, and composites reinforced with nanofibers and carbon nanotubes.
Turboelectric aircraft propulsion systems present a number of challenges related to other aircraft systems (e.g., thermal management systems). More structurally and aerodynamically efficient configurations can help address these challenges.
Aircraft turboelectric propulsion requires advances in many system capabilities. Motors, generators, and electrical distribution systems lie at the heart of electrical systems, and most electric propulsion research is understandably focused on these key elements. As these electric propulsion technologies advance, it is essential that research to advance capabilities in related aircraft systems is properly directed. Studies to date of the benefits and
challenges of turboelectric propulsion, however, have paid insufficient attention to all contributing aircraft systems. This is particularly true with regard to thermal management systems, because they are essential to the performance of the electric propulsion system and because they may affect flight performance of the aircraft. Other systems of particular interest include aircraft structure and optimized aerodynamic integration.
Research Infrastructure for Electrical Technologies
The research and development of megawatt-class turboelectric aircraft propulsion systems is hampered by the lack of development testing facilities.
Existing facilities for the development of motors, generators, and other electrical equipment for aircraft have been generally designed to provide nonpropulsive power. Electric propulsion systems will need to have much higher power capacities than electrical systems currently on aircraft. Facilities to meet these higher power levels have not been developed. No capability for proper simulation of a turboelectric propulsion system for a large aircraft exists at this time.
Turboelectric Aircraft System Studies
Conduct more encompassing studies of aircraft powered by turboelectric systems in order to better understand the benefits, component performance sensitivities, certification issues, and trade-offs related to key aircraft systems, such as thermal management and energy storage.
Establishing cost targets, thermal management targets, and reliability targets at an early stage would help define the research plan. In addition, certification plans would be most effective if established in active collaboration with the U.S. Federal Aviation Administration or other certification authorities such as the U.K. Civil Aviation Authority and the European Aviation Safety Agency.
Core Turboelectric Technologies
Develop the core technologies that are required for megawatt-class turboelectric propulsion systems: motors, generators, inverters, power distribution, and circuit protection.
It would be appropriate for research to address 1 to 5 MW systems, with an initial focus on 1 MW systems.
Megawatt-Class Research Facilities
Develop research facilities for megawatt-class electric power and thermal management systems suitable for testing turboelectric aircraft propulsion systems.
After initially focusing on ground-based facilities, a flight demonstration program for a turboelectric system could be considered in order to support advanced development of MW-class systems.