aerospace

Article Hybrid Electric Aircraft Propulsion Case Study for Skydiving Mission

Richard Glassock 1,*,†, Michael Galea 1,† ID , Warren Williams 2,† and Tibor Glesk 3

1 Institute for Aerospace Technology, The University of Nottingham, Innovation Park, Triumph Road, Nottingham NG7 2TU, UK; [email protected] 2 Engineering and Built Environment, University of Southern Queensland, West Street, Toowoomba Q4350, Australia; [email protected] 3 Air Ute Pty Ltd., Caloundra Q4551, Australia; [email protected] * Correspondence: [email protected]; Tel.: +44-0115-748-6449 † These authors contributed equally to this work.

Received: 30 June 2017; Accepted: 15 August 2017; Published: 18 August 2017

Abstract: This paper describes a case study for applying innovative architectures related to electrified propulsion for aircraft. Electric and hybrid electric propulsion for aircraft has gained widespread and significant attention over the past decade. The driver for industry interest has principally been the need to reduce emissions of combustion engine exhaust products and noise, but increasingly studies revealed potential for overall improvement in energy efficiency and mission flexibility of new aircraft types. In this work, a conceptual new type for a skydiver lift mission aircraft is examined. The opportunities which electric hybridisation offers for this role is analysed in comparison with conventional legacy type propulsion systems. For a conventional commercial skydiving mission, an all-electric propulsion system is shown as viable, and a hybrid-electric system is shown to reduce aircraft fuel costs and CO2 emissions whilst maintaining conventional aero-engine operational benefits. The new paradigm for aircraft development which hybrid electric propulsion enables has highlighted significant issues with aircraft certification practices as they exist today. The advancement of aircraft design and production to harness the value of new propulsion systems may require adaption and development of certification standards to cater for these new technologies.

Keywords: turbo-electric; hybrid; aircraft; performance; simulation; propulsion; efficiency; utility; mission; modular; configuration; certification

1. Introduction Recent advancements in aircraft propulsion technology are trending toward more electric propulsion and fully electric concepts. Over the past two decades, several key technologies for electrical power and drive systems have matured to the extent where the power and energy density has become suitable for certain aerospace applications and missions. The power density and reliability of electric motors and drive electronics have become viable, while projections on the advancement of battery storage energy density lead to industry-wide interest in aerospace electrical propulsion systems. However, studies of presently available electrical technology invariably conclude a deficit compared to traditional thermal engine propulsion for larger capacity transport missions due to the excessive battery weight. Today, there are already certain aircraft types and missions which favour electrical propulsion, but the limit is approximately two person-hours of flight time. This paper demonstrates how this same limit favours the skydiving mission under typical commercial conditions and provides an example of a new aircraft type.

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The addition of electrical propulsion systems (EPS) for aircraft functionality is not solely based on overall efficiency considerations. The term “utility” is useful to consider as defined by “the state of being useful, profitable, or beneficial” [1]. A certain minimum aircraft efficiency is clearly necessary to perform the required mission, but a trade-off can exist against adequate utility. The value of hybrid electric propulsion (HEP) must be assessed according to the combination of efficiency and utility of an aircraft design to suit a mission and a market.

1.1. Skydiving Lift Mission Skydiving is a popular aviation sport throughout the world. While the overall number of participants is relatively small compared to many other usual sports, there is a continuous popularity with hundreds of thousands of people active and approximately 1000 centres worldwide [2]. The United States Parachute Association alone recorded 36,770 members at the end of 2014 [3]. As compared to military parachuting operations, sports skydiving usually requires commercial operators or clubs to utilise converted general aviation and small commuter type aircraft to get participants to suitable jump altitude, typically up to 4000 m above ground level (AGL). The types in use for commercial operations are frequently four-seat to 10-seat aircraft, which include light aircraft and commuter category types. Examples of aircraft used include Cessna182, Cessna206 and 208, Piper PA-31, Pilatus PC-6, PC-12, LET410, and Twin Otter types, or similar piston or turbo-prop powered aircraft. As with any aircraft operation, the commercial viability of any choice is dependent on many factors such as the demand and utilisation, operating costs (maintenance, fuel), insurance etc. In many cases, the aircraft types which remain in service are aging legacy piston powered examples, typically between 20 and 50 years old, or the younger but more capital intensive turbo-props. New designs suitable for this role are warranted but there are limited options in the market place. Apart from issues of safety with older legacy types which arise due to the demanding nature of the mission and loading cycle, the overall energy utilisation and resulting emissions are significant compared to the actual time spent engaged in the primary activity, free-fall and canopy time. Whereas the sport is partly centred on natural environment experiential quality, the process is highly energy intensive. The utilisation of more renewable green energy will reduce the carbon footprint [4] and offer reduced participation costs for the sport. A new clean-sheet design could yield operational, safety and environmental benefits. Key technologies for electrical motors and power electronics systems have matured to the extent where the power and energy density has become suitable for certain aerospace applications [5,6]. Battery storage energy density is improving incrementally, but the future potential must be led by industry adoption of aerospace EPS which show viability even with modest battery performance. Studies of presently available aircraft electrical propulsion technology invariably conclude a deficit compared to traditional thermal engine propulsion for larger capacity transport missions due to excessive battery energy storage weight [7,8] and hence a limited payload-range. Therefore, little incentive exists to develop light commuter-class hybrid and electric aircraft, for existing and near-term markets. A new design utility aircraft suitable for skydiving and freight roles has been proposed by Air Ute Pty Ltd. (Caloundra, Australia), as shown in Figure1. This design features several notable design attributes which provide a compelling case for development due to extraordinary mission versatility and operational payload logistics. The skydiving role suitability is achieved according to fundamental market configuration needs, internal volume, exit door size, wing and tail layout etc. In addition, the aircraft will be highly suitable for freight and logistics roles with the capability of loading and carrying two airline transport containers. No current production aircraft in this size and category can load and carry the specified container. The aircraft itself will be able to be easily dismantled and packed into a standard “intermodal container” for ease of conventional transport by road, rail and ship, thus creating an ideal transport solution. This aircraft and its role make a very suitable baseline to consider as a candidate for a hybrid electric comparison. AerospaceAerospace2017 2017,,4 4,, 45 45 33 ofof 2222

Figure 1. Conceptual prototype skydiver and freight role aircraft by Air Ute Pty Ltd. with conventional turbopropFigure 1. propulsionConceptual system. prototype skydiver and freight role aircraft by Air Ute Pty Ltd. with conventional turboprop propulsion system. 1.2. Hybrid Electric Propulsion Suitability 1.2. Hybrid Electric Propulsion Suitability Many examples of all-electric light aircraft are emerging with approximately two person-hours of flightMany time examples endurance. of all-electric This makes light a aircraft role as are a primary emerging trainer with forapproximately example, where two person-hours the mission involvesof flight carryingtime endurance. an instructor This and makes student a role pilot as fora primary a 1 h mission, trainer feasible for example, for an EPS. where However, the mission a 1 h enduranceinvolves carrying is generally an instructor considered and inadequate student pilot when for translateda 1 hour mission, into a maximum feasible for range an EPS. for However, transport ora 1 commuter hour endurance roles. Any is generally future improvement considered inadequate in battery technologywhen translated will be into instantly a maximum adaptable range to anfor aircrafttransport already or commuter designed forroles. to integrateAny future battery improv electricement drives, in furtherbattery improvingtechnology efficiency, will be reducinginstantly costs,adaptable but also to an expanding aircraft rolealready and designed mission suitability for to integrate to new marketsbattery electric without drives, requiring further new designimproving and manufacturing.efficiency, reducing However, costs, the but viability also expanding of currently role and and immediately mission suitability foreseeable to new electrical markets propulsion without technologyrequiring new must design be shown and in manufacturing. order to justify thisHowever, new development. the viability of currently and immediately foreseeableThe energy electrical density propulsion of the battery technology must increase must be by shown at least ina order factor to of justify five to this yield new a suitabledevelopment. range commensurateThe energy with density ordinary of the present battery day must conventional increase by internal at least combustion a factor of engine five to (ICE) yield propulsion a suitable equippedrange commensurate light aircraft with [7]. ordinary While new present battery day technology conventional is envisaged internal combustion to emerge inengine the future, (ICE) improvingpropulsion the equipped range and light endurance aircraft [7]. possibilities While new while battery maintaining technology some is key envisaged advantages to emerge of the EPSin the is presentlyfuture, improving made possible the range by use and of endurance combined possibili hybrid systems.ties while An maintaining aircraft using some HEP key canadvantages utilise the of superiorthe EPS energyis presently density made ofhydrocarbon possible by use fuel of and comb theined ICE overhybrid that systems. of electrochemical An aircraft batteryusing HEP storage, can whileutilise maintaining the superior many energy of the density benefits of ofhydrocarbo superior mechanicaln fuel and efficiencythe ICE over and that reliability of electrochemical of pure EPS. battery storage, while maintaining many of the benefits of superior mechanical efficiency and 2.reliability Materials of andpure Methods EPS.

2. MaterialsAnalytic and and Methods simulation methods have been used to qualify and quantify the expected performance of the skydiver aircraft. The case study uses a numeric calculation tool based on flight mechanicsAnalytic applicable and simulation to given flightmethod conditions.s have been A parallel used studyto qualif usingy X-Planeand quantify simulation the softwareexpected wasperformance conducted of tothe corroborate skydiver aircraft. the findings, The case contrast study detailed uses a numeric configuration calculation examples tool based and provide on flight a convenientmechanics wayapplicable to compare to given expected flight conditions. performance A withparallel a well-known study using legacy X-Plane type. simulation software was conducted to corroborate the findings, contrast detailed configuration examples and provide a 2.1.convenient Mission Profileway to compare expected performance with a well-known legacy type. The role and mission profile of skydiver lift aircraft in commercial operations is known [9] 2.1. Mission Profile to favour carriage of eight jumpers (loads), to a height of 4000 m AGL with a duty cycle of 3 to 4 loadsThe per role hour. and Amission typical profile skydiving of skydiver mission lift profile aircraft is shown in commercial in Figure operations2. This implies is known a mission [9] to timefavour (or carriage endurance) of eight of jumpers between (loads), 15 and to 20 a min,height which of 4000 when m AGL given with the a duty weight cycle of of the 3 loadto 4 loads and altitudeper hour. required A typical is shownskydiving below mission to yield profile an energy is shown requirement in Figure within 2. This that implies available a mission from time current (or EPSendurance) technology. of between 15 and 20 min, which when given the weight of the load and altitude required is shownHowever, below such to yield a system an energy would requirement require a complete within that recharge, available or batteryfrom current replacement EPS technology. after each mission. This concept is practically feasible, but we consider an alternative, the inclusion of an ICE element in a hybrid scheme to ensure reserve energy availability, redundancy and reduced battery re-charge or replacement frequency to avoid operating too close to hard limits of endurance without safety margins or reserves.

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Figure 2. Typical skydiving mission profile.

However, such a system would require a complete recharge, or battery replacement after each mission. This concept is practically feasible, but we consider an alternative, the inclusion of an ICE element in a hybrid scheme to ensure reserve energy availability, redundancy and reduced battery re-charge or replacement frequencyFigure to 2.2. avoid Typical operat skydivingskydivinging too missionmission close profile.profile. to hard limits of endurance without safety margins or reserves. A standard numerical simulation technique, taking candidate power-plant and aircraft AHowever, standard such numerical a system simulationwould require technique, a complete taking recharge, candidate or battery power-plant replacement and after aircraft each aerodynamic performance parameters into account, has been used to obtain energy requirements for aerodynamicmission. This performanceconcept is practically parameters feasible, into account, but we hasconsider been usedan alternative, to obtain energy the inclusion requirements of an ICE for the given mission. By varying the ratio of installed power, and energy storage between EPS and ICE theelement given in mission. a hybrid By scheme varying to the ensure ratio reserve of installed energy power, availability, and energy redundancy storage between and reduced EPS and battery ICE drive components, certain quantities such as battery weight, climb rate, fuel consumption and drivere-charge components, or replacement certain quantitiesfrequency suchto avoid as battery operat weight,ing too climbclose to rate, hard fuel lim consumptionits of endurance and recharge without recharge rate for given technology constraints can be compared. ratesafety for margins given technology or reserves. constraints can be compared. A typical HEP concept is shown in Figure 3. The aircraft may be designed to incorporate onboard AA typicalstandard HEP numerical concept is shownsimulati inon Figure technique,3. The aircraft taking may candid be designedate power-plant to incorporate and onboard aircraft recharging, or convenient battery exchange for rapid on turn around for next mission. recharging,aerodynamic or performance convenient battery parameters exchange into account, for rapid has on turnbeen aroundused to forobtain next energy mission. requirements for the given mission. By varying the ratio of installed power, and energy storage between EPS and ICE drive components, certain quantities such as battery weight, climb rate, fuel consumption and recharge rate for given technology constraints can be compared. A typical HEP concept is shown in Figure 3. The aircraft may be designed to incorporate onboard recharging, or convenient battery exchange for rapid on turn around for next mission.

Figure 3. Figure 3. HybridHybrid electric propulsion architecture.

The hybridisationhybridisation ratioratio or or degree degree of of hybridisation hybridisatio mayn may be definedbe defined as the as ratio the ratio of the of electric the electric motor motor(EM) to (EM) ICE powerto ICE availability.power availability. A more A useful more metric useful ismetric to consider is to consider the ratio the of EMratio to of total EM installed to total installedpower, thus, power,HP =thus, PEM /PHPtot =, wherePEM/Ptot a, where fully electric a fully aircraft electric could aircraft be could considered be considered as having as a unityhavingH Pa, unityand a H fullyP, and ICE a poweredfully ICE aircraftpoweredFigure an aircraft3.H HybridP ofzero. an electric HP of propulsionzero. architecture.

2.2. The The Proposed hybridisation Models ratio or degree of hybridisation may be defined as the ratio of the electric motorCase (EM) Study to ICE 1 usespower an availability. analytical modelA more developed useful metric according is to consider to conventional the ratio ofaircraft EM to flight flight total mechanicsinstalled power, for climb thus, and H descent P = PEM/P performance. tot, where a fully Parameters electric are aircraft included could to be vary considered the electrical as having system a contributionunity HP, and forfor a energy,fully energy, ICE power poweredpower and and massaircraft mass through an through HP aof range zero. a ofrange hybridisation of hybridisation ratios. The ratios. critical The outcomes critical outcomesare the required are the electrical required system electrical mass, system including mass, battery, including and battery, climb rate and for climb the resulting rate for the gross resulting weight. 2.2. The Proposed Models grossThis yieldsweight. the This mission yields cycle the mission time, and cycle a casetime, for and viability a case for when viability compared when to compared current technology to current aircraftCase structural Study and1 uses aerodynamic an analytical limits model for basic developed and useful according load capability. to conventional The model aircraft aims to flight give qualitativemechanics andfor climb quantitative and descent data relatedperformance. to mission Parameters parameters are included against the to levelvary ofthe hybridisation. electrical system contribution for energy, power and mass through a range of hybridisation ratios. The critical outcomes are the required electrical system mass, including battery, and climb rate for the resulting gross weight. This yields the mission cycle time, and a case for viability when compared to current

Aerospace 2017, 4, 45 5 of 22 technology aircraft structural and aerodynamic limits for basic and useful load capability. The model aims to give qualitative and quantitative data related to mission parameters against the level of hybridisation. AerospaceThe2017 fundamental, 4, 45 equation used to determine the climb rate, VC for relatively small thrust-to-5 of 22 weight ratios is; − The fundamental equation used to determine the climb rate, VC for relatively small = (1) thrust-to-weight ratios is; − where PA = power available, PR = power required. IfPA detailedPR knowledge of thrust available at various VC = (1) climb speeds and altitudes is known, then PA is calculatedW as the product of true airspeed and the thrustwhere componentPA = power along available, the flightPR = powerpath. Alternatively, required. If detailed an estimate knowledge for PA ofcan thrust be made available as the at product various ofclimb propeller speeds shaft and power altitudes and is th known,e propulsive then P Aefficiencyis calculated [10]. asPR themay product be estimated of true by airspeed the following and the formulathrust component [11]; along the flight path. Alternatively, an estimate for PA can be made as the product of propeller shaft power and the propulsive efficiency/ [10]. PR may be estimated by the following 4√2 (/) formula [11]; _ = / (3√) 1/4 s 4 2C (W/S ) P = D0 W W where estimates are required forR C_minimumD0 the zero lift drag co-efficient,3/4 and , the Oswald efficiency factor. (3πeRA) ρ RA, the wing aspect ratio taken as 10 and SW the wing area, 20 m2 are given from the basic aircraft geometrywhere estimates while are the required ambient for airC densityD0 the zero varies lift with drag flight co-efficient, altitude and accordinge, the Oswald to a standard efficiency function. factor. 2 ForRA, the the purposes wing aspect of this ratio analysis, taken asCD0 10 assumed and SW asthe 0.025 wing and area, equal 20 m toare 0.835. given The from significant the basic variables aircraft whengeometry comparing while ρ acrossthe ambient different air propulsi density varieson installations with flight are altitude therefore according only W to, the a standard weight which function. is theFor sum the purposes of the empty of this weight analysis, (includingCD0 assumed battery as0.025 and andnominale equal fuel) to 0.835.plus payload, The significant and propeller variables efficiency.when comparing Using acrossthis technique, different propulsionthe direct installationseffects of the are propulsion therefore only systemW, the mass weight and which propeller is the efficiencysum of the can empty be isolated weight and (including examined. battery and nominal fuel) plus payload, and propeller efficiency. UsingFigure this technique,4 illustrates the a conceptual direct effects prototype of the propulsion “Skybrid” system hybrid mass electric and skydiving propeller aircraft efficiency based can on be theisolated Air Ute and development. examined. In comparison to Figure4 illustrates a conceptual prototype “Skybrid” hybrid electric skydiving aircraft based on the AirFigure Ute 1 development., the nose mounted In comparison turboprop power-plant to Figure1, the is augmented nose mounted by two turboprop wing-mounted power-plant electric is motoraugmented and propeller by two wing-mountedinstallations. This electric propeller motor layout and propeller is effectively installations. the same Thisas the propeller rather simple layout andis effectively well established the same “tri-motor” as the rather seen simple in many and well types established historically; “tri-motor” however, seen new in manysignificant types advantageshistorically; of however, this distributed new significant propulsion advantages system ofusing this a distributed hybrid electric propulsion scheme system are available using a which hybrid ordinaryelectric scheme legacy aremulti-engine available whichlayouts ordinary cannot offer. legacy multi-engine layouts cannot offer.

FigureFigure 4. 4. ConceptualConceptual prototype prototype “Skybrid” “Skybrid” hybrid hybrid electric electric skydiving/freight skydiving/freight aircraft. aircraft.

2.3.2.3. Mission Mission Energy Energy TheThe overall overall mission mission energy energy required required is is calculated calculated from from the following items: •• PotentialPotential energy energy requirements requirements to to climb; climb; •• DragDrag energy energy expended expended in in climb; climb; •• EnergyEnergy regeneration regeneration during during descent. descent.

ItIt is notednoted that that the the drag drag energy energy expended expended during during the descent the descent phase ofphase the flight of the can flight be considered can be consideredto be fully regeneratedto be fully regenerated from the potential from the energy potential gained energy during gained climb during (a pure climb glide), (a pure and glide), is a form and of isregeneration. a form of regeneration. However, forHowever, operational for operationa reasons, forl reasons, non-hybrid, for non-hybrid, or very low or proportionsvery low proportions of electric ofpropulsion, electric propulsion, the ICE must the beICE running must be during running the descent.during the The descent. aircraft mustThe aircraft maintain must the maintain capability the for re-routed flight profiles such as dictated by air traffic control, or events such as go-arounds, holds, or diversions. An ICE unit takes a certain amount of time to start and warm-up; therefore, it is typical and usually mandatory for the ICE to be running even during a full glide descent profile. The thermal efficiency of the ICE at flight idle is typically lower than at climb or cruise conditions. Aerospace 2017, 4, 45 6 of 22

Regeneration for hybrid electric propulsion can include energy derived from using the propeller as an air turbine during unpowered descent. Although the efficiency of the “wind-milling” operation is low due to the incorrect shape of the propeller as a turbine, and due to the Betz limit, it has been shown that between 5% and 40% [7,12] of the potential energy can be regenerated in this way. The resulting battery recharge quantity is very low, since much of the climb energy expenditure is not toward gravitation potential, but is lost as drag. However, the small gain comes with added benefits. For the skydiving operation, a rapid descent is necessary to allow improved cycle time. Ordinary aircraft descent rates are limited by the maximum permissible airspeed and deployable drag. Extra drag to allow steeper descent is usually afforded by the use of wing-flaps. Utilising the aircraft propeller as a turbine implies very high drag, this can enable very steep descent within certificated airspeed limitations. Propulsion efficiency for propeller driven aircraft is determined by the design airspeed, thrust requirements and the propeller design characteristics. The primary design variables are the propeller pitch and diameter, and both must also suit the characteristics of the power-plant. Current technology aircraft have a very well-established design synthesis which can often yield a propulsive efficiency above 85% for specific flight conditions. In this case, there is not very much improvement available, but the implementation of hybrid electric components should aim to allow improvement to either the range of high efficiency throughout the flight profile, or a reduction in the complexity required to obtain the peak efficiency. For example, the addition of the EPS to the ICE can allow torque application characteristics which can eliminate the need for variable pitch propeller mechanisms, or allow larger diameter propellers without gear boxes, more suitable for high efficiency lower speeds [12]. Utilisation of EPS can provide more overall propeller disk area for a particular aircraft configuration, and thus improve propulsive efficiencies, particularly at lower speeds. Adding extra propeller disk area may be accomplished by adding EPS units, far more easily than adding extra ICE units, as shown by the concept illustrated in Figure4. The mass specific power of the EM is presently comparable to the best ICE and is forecast to further improve in the near term while being orders of magnitude simpler in terms of parts count, manufacturability and a range of other attributes. The EM is extremely easy to mount on the airframe without significant structural or systems weight and complexity, these qualities are enabling many new aircraft designs both in concept and practice today. The provision of increased thrust, at lower airspeeds for a given aircraft mass implies that steeper climb angles are possible. In general, a steeper climb angle, at a lower airspeed can result in significantly reduced drag for a given climb rate. In particular, the induced drag decreases with the climb angle, and the induced drag is typically triple the value of parasite drag in climb configurations [11]. Aerodynamic efficiency improvements are also feasible by taking advantage of carefully integrated thrust characteristics which can reduce drag. Often a distributed propulsion (DP) scheme is used to provide powered lift capability in order to reduce the required wing area [13], or may be selected to ingest the turbulent boundary layer in order to reduce overall drag [14]. In this analysis, such advantages are not considered in depth, but remain as further potential for a conservative approach. Another item of propeller efficiency is related to the slipstream effect on the airframe components according to the location of the propeller. This appears to have a very significant influence and is discussed in further in this work.

2.4. ICE Limitations For the skydiving mission, the maximum rate of descent for the aircraft is an important parameter, as it directly affects the mission cycle time. Several aspects of ICE installation can limit the maximum descent rate. The natural wind-milling characteristic of propellers can be a disadvantage. Depending on the ICE in use, piston or turbine, and the propeller type, fixed or variable pitch, the engine can be damaged if over-driven by the wind-milling propeller. This damage must be avoided by limiting the descent rate and airspeed. Aerospace 2017, 4, 45 7 of 22

Another limiting factor for conventional ICE powered aircraft descent is the behaviour of the ICE cooling system. Air cooled piston engines of the type very often used in current skydiving operations are particularly susceptible to shock cooling, caused by transition from high power settings at low airspeed to low power at high airspeed. This is the typical situation for aircraft in this role. Turboprop powered aircraft are less susceptible to shock cooling, but nonetheless, strict attention to temperature limits and variation must be observed.

2.5. Rated ICE Power Internal combustion engine maximum output power is typically limited by any of three stress factors: pressure, engine speed (RPM) and temperature. The internal gas pressure developed during the operating cycle is dependent on the ambient atmospheric pressure but can be significantly modified by the nature of the engine mechanical systems. A maximum rating will usually be limited to a particular continuous operating time, and is typically given for a standard day sea level temperature and pressure. A “naturally aspirated” piston engine’s maximum power output varies with altitude due to the change in air density and therefore the amount of fuel which can be burned. A “forced induction” piston engine is fitted with a compressor to increase the density of the incoming air and is able to maintain full-rated power at higher altitudes. Gas turbine engines always feature a compressor and also have the potential to maintain full-rated power at higher altitudes. Nevertheless, because the rated power of an engine is always limited by maximum stress, a compressor equipped type may need to be operated below its theoretical maximum power when at lower altitudes. The resultant power rating is termed “flat”, when the available maximum engine power is constant up to some altitude. In this paper, a flat-rated engine is considered as having the same performance as a full-rated unit (at sea level) except that it maintains maximum rated power at altitude. Whereas the full-rated unit is assumed to lose power proportionally by the ratio of air density at altitude versus sea level. It is a notable attribute of EPS power-plants that they are not subject to output variation due to altitude, although thermal and RPM limits are still applicable.

2.6. Recharge Cases Two particular HEP cases are analyzed and discussed in this work. This comparison shows advantages and disadvantages according to how the battery charging of a HEP aircraft should be managed. Case A considers a hybrid electric aircraft propulsion system configured with a battery which is recharged using the on-board engine. Therefore, the aircraft can be fully self-contained and reliant only on consumable fuel, always maintaining a fully charged battery at the end of the mission. In this strategy, the ICE and EPS would provide power for flight but the duty cycles would be specified such that the ICE could always be capable of recharging the battery as required during the mission. An inevitable consequence of this strategy is that the electrical energy comes at the cost of the ICE efficiency, which is seldom better that 25%. Another problem with this approach is that the generator system capacity must be sized according to the engine as well as a time allocation based on the descent time. In the skydiving role, the descent time must be minimized to ensure high mission rates, and the electrical energy used from the battery during climb must be replaced during descent. The intent is to reduce the size and power of the ICE required for the mission, while retaining the capability for ICE-only operations independent of EPS storage. In addition, the cost and complexity of ICE units strongly favour the use of a single engine, and the accompanying EPS when in generator mode must then be sized to take the full ICE power. For a single ICE configuration, at higher HP levels the power applied to non-ICE EPS units in climb will be significantly more than that able to be regenerated during descent. This adversely affects the recharge time which should be minimized as necessary to complete within the prescribed descent cycle time. Aerospace 2017, 4, 45 8 of 22

An inherent complication with “Case A” configuration is the sizing of the EPS equipment for efficient utilisation of ICE electrical generation. Clearly, a low HP aircraft will have a lot of excess ICE power to recharge on descent, but also a low EPS capacity, whereas a high HP aircraft would have a low ICE power available but a high EPS charging capacity and requirement. These factors will influence the viability of the “Case A” type for this mission. Case B considers a hybrid electric aircraft propulsion system configured with a battery which is recharged by ground based sources. The battery can be removed and refitted easily such that the aircraft can resume operation for the next mission using a pre-charged exchange battery. It is shown that Case B is a superior system in terms of energy efficiency, direct operating costs and emissions, but may be significantly restricted in operational versatility. Nevertheless, the Case B system is presently feasible given current EPS technology, and will be ready for any and all improvements in battery storage technology over time which only improve upon original performance and utility.

3. Results

3.1. Hybrid Electric Modelling has yielded results as presented later in this section which show that the full spectrum of hybridisation ratio is feasible right through to full electric operation for this aircraft type and mission. The modelling for Case A indicates that the overall fuel consumption will increase if an on-board powerplant is used to recharge the battery during the operating cycle, using a conventional gas turbine unit. The specific fuel consumption for the preferred candidate type of ICE available, a gas turbine, turbo-prop or turbo-shaft unit is typically around 0.3 kg per kilowatt-hour [8]. While this type of engine has significant advantages in terms of mass specific power, approximately 4 kW per kilogram, and is a mature propulsion system within aerospace applications, significantly higher thermal efficiency would be required to make it viable for on-board recharging compared to ground based power sources. The application of electrical power to the aircraft, necessitates carriage of extra weight for batteries and power electronics, therefore the energy required for the flight is increased, unless efficiency is gained in either propulsion or aerodynamic efficiency. Nevertheless, the capability of configuring the aircraft for different missions using the EPS at certain HP levels could find utility for some operators. The modelling for Case B, indicates that HEP has the potential to significantly reduce emissions and direct operating costs for the flight operations.

3.2. Model Assumptions The analytic model is based on standard flight mechanics run as a MATLAB tool. Without highly detailed analysis and wind-tunnel testing, certain assumptions on aerodynamic characteristics are required as input variables. The primary aerodynamic variable which must be estimated for this model is known as the Oswald efficiency factor (e). This is a parameter which expresses the total variation of drag with lift. It is sometimes called the span efficiency factor and would equal 1.0 for an elliptically-loaded wing with no lift-dependent viscous drag, for practical aircraft “e” varies from about 0.75 to 0.90 [15]. The other main variable which must be assumed for the model is the propeller efficiency. Again, values between 0.75 and 0.90 are typical [11] for the type of aircraft under consideration. Also, as noted below, the best propeller efficiency may be considered as dependent on the detailed aircraft configuration, with wing mounted propellers usually resulting in greater efficiency than a front fuselage mounted unit due to reduced slipstream “scrubbing”. For the results below, an “e” value of 0.835 was used, and propeller efficiency varied between 0.75 and 0.85 respectively for the non-hybrid through to fully electric models assuming any EPS is accomplished by adding wing-mounted propellers. For intermediate HP, the propeller efficiency was linearly scaled. Increasing HP in this analysis implies increasing both the propeller area and the proportion assumed to be wing mounted and unobstructed by the fuselage. The battery mass required is calculated by a numerical iterative process taking into account the energy requirement according to the time required for climb Aerospace 2017, 4, 45 9 of 22 Aerospace 2017, 4, 45 9 of 22 assumed to be wing mounted and unobstructed by the fuselage. The battery mass required is calculated by a numerical iterative process taking into account the energy requirement according to atthe changing time required weight and for drag climb conditions. at changing The overallweight EPSand massdrag isconditions. calculated usingThe overall the proportion EPS mass of is installedcalculated power using multiplied the proportion by the power of installed density powe andr ismultiplied added to batteryby the power mass. Thedensity same and technique is added is to used to calculate the ICE mass in proportion to H to arrive at a take-off weight given the addition of battery mass. The same technique is used to calculateP the ICE mass in proportion to HP to arrive at a sufficienttake-off mission weight fuel.given Table the 1addition, power systemof sufficient analytic mission model fuel. parameters, Table 1, andpower Table system2, aircraft analytic analytic model modelparameters, parameters, and showTable the2, aircraft key parameters analytic usedmodel in parameters, the analytical show calculations. the key parameters used in the analytical calculations. Table 1. Power system analytic model parameters.

Table 1. Power system analytic model parameters. Component Energy Density (Wh/kg) Power Density (kW/kg) Efficiency EMComponent Energy Density− (Wh/kg) Power Density6 (kW/kg) Efficiency 0.9 − Power ElectronicsEM − 66 0.9 0.9 Battery 200 − 0.9 ICEPower Electronics − 46 0.9 0.26 Battery 200 − 0.9 ICE Table 2. Aircraft analytic model parameters. 4 0.26

Installed TableEmpty 2. Aircraft Weight analytic Payloadmodel parameters. Oswald Propeller Aircraft Model Power (kW) 1 (kg) Weight 1 (kg) Efficiency Factor Effciency Installed Empty Payload Oswald Propeller AUXXAircraft Model 670Power (kW) Weight 2700 1 (kg) Weight 800 1 (kg) Efficiency 0.835 Factor Effciency 0.75 Conventional AUXX AUXX Hybrid 670670 2700 + %EPS2700 800 800 0.835 0.835 Scaled 0.75 AUXXConventional 670 3500 800 0.835 0.85 All-ElectricAUXX Hybrid 670 2700 + %EPS 800 0.835 Scaled AUXX All-Electric 670 3500 800 0.835 0.85 1 Weight given as mass in kilograms including nominal mission fuel. 1 Weight given as mass in kilograms including nominal mission fuel. FigureFigure5 shows 5 shows the the recharge recharge time time over-run over-run for afor “Case a “Case A” A” hybrid hybrid mission mission using using a single a single ICE/EPS ICE/EPS unitunit (single (single propeller). propeller). The The “time “time over-run” over-run” is the is amountthe amount of time of time the enginethe engine driven driven generator generator would would be requiredbe required to continue to continue running running on the groundon the afterground landing after tolanding recharge to therecharge battery. the As battery. the HP increases, As the HP moreincreases, battery more energy battery is consumed energy is forconsumed the climb, for butthe climb, less ICE but power less ICE is availablepower is available for both thefor both climb the andclimb the descentand the phase.descent The phase. practical The practicalHP limit H forP limit Case for A mustCase lieA must somewhere lie somewhere before the before knee the of the knee curve,of the around curve,H Paround= 0.8. However,HP = 0.8. atHowever, the value atH theP = value 0.5, the H over-runP = 0.5, the is10%, over-run which is would 10%, which represent would a reasonablerepresent limit a reasonable considering limit the considering ground handling the ground and operationalhandling and margins. operational margins.

FigureFigure 5. Hybridisation5. Hybridisation proportion proportion vs. vs. Case Case A chargeA charge time time over-run; over-run; single single EPS. EPS.

If multiple EPS units are used (three in this example), with a single ICE, the problem of recharge If multiple EPS units are used (three in this example), with a single ICE, the problem of recharge time becomes even more severe; as shown in Figure 6, where the useful HP value is reduced to 0.25. time becomes even more severe; as shown in Figure6, where the useful HP value is reduced to 0.25.

Aerospace 2017, 4, 45 10 of 22 Aerospace 2017, 4, 45 10 of 22 Aerospace 2017, 4, 45 10 of 22

Figure 6. Hybridisation proportion vs. Case A charge time over-run; multiple EPS. FigureFigure 6. Hybridisation 6. Hybridisation proportion proportion vs. vs. Case Case A chargeA charge time time over-run; over-run; multiple multiple EPS. EPS. The problems inherent in Case A strategy in terms of mission cycle time are minor compared to The problems inherent in Case A strategy in terms of mission cycle time are minor compared to theThe significant problems increase inherent overall in Case energy A strategy consumpt in termsion. Figure of mission 7 shows cycle the time mission are minor fuel compared recharge the significant increase overall energy consumption. Figure 7 shows the mission fuel recharge torequirements the significant increase increase with overall increasing energy hybridisation consumption. ratio. Figure That is7 additionalshows the recharge mission fuel fuel is recharge required to fullyrequirements recharge increase batteries with during increasing a typical hybridisation skydiving ratio.mission, That based is additional on a mission recharge of durationfuel is required 15–20 requirementsto fully recharge increase batteries with increasing during a hybridisationtypical skydiving ratio. mission, That is based additional on a mission recharge of fuelduration is required 15–20 min. to fullymin. recharge batteries during a typical skydiving mission, based on a mission of duration 15–20 min.

Hybridization versus Mission Fuel Case A On-board Recharge 70 Hybridization versus Mission Fuel Case A On-board Recharge 70 Mission Fuel required Flat Rated 65 MissionMission Fuel Fuel required required Full Flat Rated Rated 65 Mission Fuel required Full Rated 60 60 55 55 50 50 45 45 40 40 35 35

30 30

25 25

20 20 00 0.1 0.1 0.2 0.2 0.3 0.3 0.4 0.4 0.5 0.5 0.6 0.6 0.7 0.7 0.8 0.8 0.9 0.9 1 1 HybridizationHybridization Proportion Proportion Electric Electric FigureFigureFigure 7. 7.On-board On-board7. On-board ICE ICE ICE recharge recharge recharge mission mission mission fuel fuel fuel required, re required,quired, assuming assuming assuming nono no chargecharge charge timetime time limit. limit.limit.

In thisIn this case, case, a non-hybrid a non-hybrid ICE-only ICE-only system system would would use use about about 25 25 kg kg of of fuel fuel to to complete complete thethe mission.mission. In this case, a non-hybrid ICE-only system would use about 25 kg of fuel to complete the mission. ForFor a flat-rated a flat-rated engine engine at atthe the H PH valueP value of of 0.25, 0.25, which which as as noted noted above above would would be be the the maximum maximum forfor viable For a flat-rated engine at the H value of 0.25, which as noted above would be the maximum for missionmission time time utilisation, utilisation, the the missionP mission fuel fuel is isincr increasedeased by by 40%. 40%. Clearly, Clearly, thethe “Case“Case A”A” strategystrategy is viable mission time utilisation, the mission fuel is increased by 40%. Clearly, the “Case A” strategy is counterproductivecounterproductive from from a fuel a fuel consumption consumption perspectiv perspective,e, and and shows shows why, why, historically, historically, hybridhybrid electric counterproductive from a fuel consumption perspective, and shows why, historically, hybrid electric aircraftaircraft concepts concepts were were not not considered considered useful. useful. Howe However,ver, there there are are certain certain advantages advantages whichwhich must be aircraft concepts were not considered useful. However, there are certain advantages which must considered,considered, and and tend tend to tofavour favour hybridisation, hybridisation, such such as as those those shown shown in in thethe followingfollowing sections.sections. The be considered, and tend to favour hybridisation, such as those shown in the following sections. justificationjustification of hybridof hybrid electric electric aircraft aircraft from from the the fuel fuel consumption consumption perspective perspective is is heavily heavily dependentdependent on Theany justificationany electric electric machine, ofmachine, hybrid propulsive propulsive electric aircraft and and aerodynamic aerodynamic from the fuel ef efficiency consumptionficiency gains gains which perspective which can can be be is mademade heavily asas dependentwellwell as the onenergy anyenergy electric density density machine, of ofbattery battery propulsive storage storage andsystems. systems. aerodynamic Each Each of of efficiencythese these factors factors gains areare which shownshown can inin be literatureliterature made as to well be asadvancing theadvancing energy and density and have have ofa positive a battery positive development storage development systems. future future Each [16]. [16]. of these factors are shown in literature to be advancingTheThe andCase Case have B typeB atype positive can can include include development on-board on-board futurerecharging recharging [16]. if if desired, desired, but but the the baseline baseline isis toto replacereplace and groundTheground Caserecharge recharge B type the the canbattery battery include each each on-board mission. mission. rechargingWhen When battery battery if desired,recharging recharging but is is the conducted conducted baseline viavia is to groundground replace based and groundsources,sources, recharge renewable renewable the batteryand and green green each energy, mission.energy, with with When many many battery wider wider recharging societal societal and and is conducted technicaltechnical advantagesadvantages via ground can based be sources,utilised.utilised. renewable This This system system and guar greenguaranteesantees energy, the the best with best emissions manyemissions wider reduction reduction societal an an anddd provides provides technical thethe advantages bestbest futurefuture canutility be potential as battery storage limits increase in the future. utilised.potential This as battery system storage guarantees limits the increase best emissions in the future. reduction and provides the best future utility potential as battery storage limits increase in the future.

Aerospace 2017, 4, 45 11 of 22 Aerospace 2017, 4, 45 11 of 22

IfIf the the presumption presumption is is made made for for Case Case B, B, where where the the battery battery is is replaced replaced or or recharged recharged without without using using thethe on-board on-board ICE, ICE, the the mission mission fuel fuel required required naturally naturally tends tends to to zero zero when when the the aircraft aircraft is is solely solely electric electric poweredpowered ( H(HP P= =1), 1),as asshown shownin inFigure Figure8 .8.

FigureFigure 8. 8.Skydive Skydive mission mission fuel fuel required required assuming assuming off-board off-board recharge. recharge.

The fuel required is nearly in proportion to the hybridisation for the flat-rated engine. In The fuel required is nearly in proportion to the hybridisation for the flat-rated engine. In addition, addition, clearly notable from Figure 8 is the difference between the flat-rated and full-rated engine clearly notable from Figure8 is the difference between the flat-rated and full-rated engine type. type. The principle reason for the extra fuel required for the full-rated engine is the extra time The principle reason for the extra fuel required for the full-rated engine is the extra time required for required for the climb. This effect diminishes with increased HP but would make hybridisation more the climb. This effect diminishes with increased H but would make hybridisation more favourable if favourable if use of a full-rated engine was necessaryP for any reason. The choice of hybridisation use of a full-rated engine was necessary for any reason. The choice of hybridisation proportion can be proportion can be made according to preferences for the energy source and hence emission made according to preferences for the energy source and hence emission characteristics, operating characteristics, operating costs, weight or other operational characteristics. The fuel required reflects costs, weight or other operational characteristics. The fuel required reflects the direct operating cost the direct operating cost of the mission as well as CO2 and NOx emissions. The overall energy required of the mission as well as CO2 and NOx emissions. The overall energy required for the mission may for the mission may increase for higher HP because of the inherent electric energy source, distribution, increase for higher H because of the inherent electric energy source, distribution, conversion and conversion and storageP inefficiencies. Carefully selecting the nature of the energy supply can easily storage inefficiencies. Carefully selecting the nature of the energy supply can easily reduce the cost and reduce the cost and environmental impact. For example, converting the energy in coal to electricity environmental impact. For example, converting the energy in coal to electricity to charge batteries via to charge batteries via a conventional national grid for flight purposes is counterproductive to a conventional national grid for flight purposes is counterproductive to efficiency and emissions goals, efficiency and emissions goals, but utilizing solar, wind or even nuclear energy sources can reduce but utilizing solar, wind or even nuclear energy sources can reduce the emissions. Nevertheless, the the emissions. Nevertheless, the EPS system weight for a given flight will exceed that required for EPS system weight for a given flight will exceed that required for pure ICE operation, this translates to pure ICE operation, this translates to more energy being required to account for the higher drag due more energy being required to account for the higher drag due to increased required wing lift and to increased required wing lift and more work done to raise the mass against gravity during the more work done to raise the mass against gravity during the climb. climb. The Case B fuel consumption naturally favours this hybridisation model; however, the viability of The Case B fuel consumption naturally favours this hybridisation model; however, the viability using such an amount of electrical storage in terms of the aircraft structural and performance capability of using such an amount of electrical storage in terms of the aircraft structural and performance must be verified. capability must be verified. At the maximum HP, the battery weight including a 20% margin amounts to 800 kg. This battery At the maximum HP, the battery weight including a 20% margin amounts to 800 kg. This battery weight is considered feasible for an aircraft with a take-off weight of 4500 kg and a payload of 800 kg. weight is considered feasible for an aircraft with a take-off weight of 4500 kg and a payload of 800 Figure9 shows the battery weight implication for both the full-rated and flat-rated hybrid electric kg. scenarios, while Figure 10 shows the effect on take-off weight. Figure 9 shows the battery weight implication for both the full-rated and flat-rated hybrid The mission cycle time is naturally dependent on the best climb rate achievable, as well as the electric scenarios, while Figure 10 shows the effect on take-off weight. descent rate. For an identical installed propulsive power of 670 kW, the analytic model shows climb The mission cycle time is naturally dependent on the best climb rate achievable, as well as the times for the range of H given the full-rated and flat-rated hybrid electric scenarios shown in Figure 11. descent rate. For an identicalP installed propulsive power of 670 kW, the analytic model shows climb times for the range of HP given the full-rated and flat-rated hybrid electric scenarios shown in Figure 11.

Aerospace 2017, 4, 45 12 of 22 Aerospace 2017, 4, 45 12 of 22 Aerospace 2017,, 4,, 4545 12 of 22 Hybridization versus Battery Mass 800 Hybridization versus BatteryBattery Mass Mass Flat Rated 800 Hybridization versus Battery Mass 700800 Battery Mass Full Rated Battery Mass Flat Rated 700 Battery Mass Full Rated 600700

600600 500

500500 400

400400 300

300300 200

200200 100 100100 Eta = 0.75 Eta = 0.85 p p 0 Eta = 0.75 Eta = 0.85 Etap = 0.75 Etap = 0.85 0 0.1p 0.2 0.3 0.4 0.5 0.6 0.7p 0.8 0.9 1 00 00 0.1 0.1 0.2 0.2Hybridization 0.3 0.3 0.4 0.4 Proportion 0.5 0.5 0.6 0.6 Electric 0.7 0.7 0.8 0.8 0.9 0.9 1 1 Hybridization Proportion Electric Figure 9. Battery mass requirements. FigureFigure 9.9. BatteryBatteryBattery mass mass requirements.requirements.

Hybridization versus Take-off Weight 4500 Hybridization versus Take-off Weight 45004500 Take-off Weight Flat Rated Take-off Weight Flat Rated 4400 Take-offTake-off Weight Weight Flat Full Rated Rated 4400 4400 Take-offTake-off WeightWeight FullFull RatedRated 4300 43004300

4200 42004200

4100 41004100

400040004000

390039003900

380038003800 Eta = 0.75 EtaEtap== 0.75 0.75 370037003700 pp Eta = 0.85 EtaEtap == 0.85 0.85 pp 36003600 3600 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 00 0.1 0.1 0.2 0.2 0.3 0.3 0.4 0.5 0.6 0.6 0.7 0.7 0.8 0.8 0.9 0.9 1 1 Hybridization Proportion Electric HybridizationHybridization Proportion Electric Electric Figure 10. Take-offTake-off grossgross weightweight vs.vs. hybridhybrid proportion.proportion. Figure 10. Take-off gross weightweight vs.vs. hybrid proportion.

Hybridization versus Climb Time 1414 Hybridization versus Climb Time 14 Climb Time Flat Rated 13 Climb Time Flat Rated 13 ClimbClimb Time Time Full Flat Rated Rated 13 Climb Time Full Rated 1212 12 1111 11 1010 10 99 9 88 8 77 7 66 Eta = 0.75 Etap = 0.75 6 p Eta = 0.85 5 Etap = 0.85 5 p Eta = 0.75 5 p Eta = 0.85 44 p 00 0.1 0.1 0.2 0.2 0.3 0.3 0.4 0.4 0.5 0.5 0.6 0.6 0.7 0.7 0.8 0.8 0.9 0.9 1 1 4 Hybridization Proportion Electric 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 Hybridization Proportion Electric Figure 11. TimeTime toto climbclimb vs.vs. hybridhybrid proportiproportion for flat-rated and full-rated ICE. Figure 11. Time to climb vs. hybrid proportionproportion for flat-ratedflat-rated and full-rated ICE.ICE. The time to climb for a fully electric example isis acceptable,acceptable, andand thethe improvementimprovement overover aa full-full- ratedThe powerplant time to climb operating for a fullyto the electricprescribed example altitude is isacceptable, very significant. and the improvement over a full- TheThe time effect to climbof propeller for a fully efficiency electric (Eta examplep) on the is climb acceptable, performance and the in improvement this model is very over sensitive. a full-rated rated powerplantThe effect of operating propeller toefficiency the prescribed (Etap) on altitude the climb is veryperformance significant. in this model is very sensitive. powerplantThe foregoing operating examples to the used prescribed Etap values altitude linearly is scale very from significant. 0.75 to 0.85 on the HP domain and show TheThe foregoing effect of examples propeller used efficiency Etap values (Eta plinearly) on the scale climb from performance 0.75 to 0.85 in on this the model HP domain is very and sensitive. show adequateThe effect performance. of propeller If efficiencythe propeller (Eta efficiencyp) on the isis climb heldheldperformance constantconstant (0.75(0.75 in forfor this example),example), model is andand very allall sensitive. otherother The foregoing examples used Etap values linearly scale from 0.75 to 0.85 on the HP domain and show The foregoing examples used Eta values linearly scale from 0.75 to 0.85 on the H domain and show adequate performance. If the propellerp efficiency is held constant (0.75 for example),P and all other adequate performance. If the propeller efficiency is held constant (0.75 for example), and all other

Aerospace 2017, 4, 45 13 of 22

Aerospacevariables2017, 4remain, 45 the same, the following plot of hybridisation versus climb time is produced shown13 of 22 Aerospacein Figure 2017 12., 4 , 45 13 of 22 In this case, the climb rate has decreased significantly with increasing HP and would have variablesvariablesnegative remain implicationsremain the the same, same, for theaircraft the following following utility. plotThis plot ofin ofdicates hybridisation hybridisation the extreme versus versus importance climb climb time time of isimproving is produced produced items shownshown of in Figureinefficiency Figure 12. 12. in order to enable hybrid electric propulsion systems to be viable. In this case, the climb rate has decreased significantly with increasing HP and would have Hybridization versus Climb Time negative implications for aircraft14 utility. This indicates the extreme importance of improving items of efficiency in order to enable hybrid electric propulsion systemsClimb Time to Flat be Rated viable. 13 Climb Time Full Rated

12 Hybridization versus Climb Time 1114 Climb Time Flat Rated 1013 Climb Time Full Rated

129

118

107

69 Eta = 0.75 p Eta = 0.75 58 p

47 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 6 Hybridization Proportion Electric Eta = 0.75 5 p Eta = 0.75 Figure 12. Time to climb vs. hybrid proportion given constantp propeller efficiency. Figure 12. Time to climb vs. hybrid proportion given constant propeller efficiency. 4 Results obtained using the0 X-Plane 0.1 0.2 simulation 0.3 0.4 0.5 as 0.6 detailed 0.7 0.8 in 0.9the following 1 section show a Hybridization Proportion Electric definiteIn this improvement case, the climb in rate rate has of climb decreased for the significantly hybrid models with compared increasing toH conventionalP and would propulsion. have negative Figure 12. Time to climb vs. hybrid proportion given constant propeller efficiency. implicationsThe X-Plane for models aircraft were utility. configured This indicates at a constant the extreme HP of 0.67. importance The analytical of improving study shown items above of efficiency does in ordernot reflect to enable such an hybrid improvement electric propulsion given the assumed systems prop to beeller viable. efficiency parameters, so it is of interest Results obtained using the X-Plane simulation as detailed in the following section show a toResults see what obtained change of using propeller the X-Plane efficiency simulation would yield as detailed results which in the concur following with section the simulation. show a definiteThe definite improvement in rate of climb for the hybrid models compared to conventional propulsion. improvementoriginal propeller in rate ofefficiencies climb for used the hybrid in the modelsabove ca comparedlculations tovaried conventional proportionally propulsion. from 0.75 The for X-Plane a The X-Plane models were configured at a constant HP of 0.67. The analytical study shown above does modelsconventional were configured non-hybrid at using a constant a singleH propellerof 0.67. thro Theugh analytical to 0.85 for study a fully shown hybrid above propulsion does notsystem reflect not reflect such an improvement given theP assumed propeller efficiency parameters, so it is of interest suchusing an improvement two additional given wing the mounted assumed propellerpropellers. efficiency These numbers parameters, are soarbitrary it is of interestand reasonable to see what toassumptions, see what change but a of much propeller more efficiency rigorous wouldand comprehensive yield results which study concurof particular with the detailed simulation. designs The change of propeller efficiency would yield results which concur with the simulation. The original originalwould be propeller necessary efficiencies to improve used estimates, in the orabove the fligcalculationsht testing ofvaried prototypes proportionally to verify realfrom outcomes 0.75 for a propeller efficiencies used in the above calculations varied proportionally from 0.75 for a conventional conventionalwould be required. non-hybrid However, using for a singlethe purposes propeller of thisthro study,ugh to a 0.85 new for set a of fully possible hybrid propeller propulsion efficiency system non-hybrid using a single propeller through to 0.85 for a fully hybrid propulsion system using two usingnumbers two can additional be conveniently wing mountedinput and propellers.the resulting These performance numbers calculated are arbitrary to find anda match reasonable to the additionalassumptions,simulation. wing Figure but mounted a 13 much shows propellers. more the riresultgorous These using and numbers a comprehensivemore areextreme arbitrary improvement study and of reasonableparticular of propeller detailed assumptions, efficiency designs but a much more rigorous and comprehensive study of particular detailed designs would be necessary wouldterms through be necessary the domain to improve of HP. estimates,Figure 13 shows or the that flig htthe testing climb time of prototypes has reduced to byverify a similar real outcomesamount to improvewouldas the simulationbe estimates,required. tool However, or predicts. the flight for the testing purposes of prototypes of this study, to verifya new set real of outcomes possible propeller would be efficiency required. However,numbers for can the be purposes conveniently of this input study, and athe new resulting set of performance possible propeller calculated efficiency to find numbersa match to can the be convenientlysimulation. input Figure and 13 the shows resulting the result performance using a calculatedmore extreme to find improvement a match to theof propeller simulation. efficiency Figure 13 showsterms the through result using the domain a more of extreme HP. Figure improvement 13 shows that of propellerthe climb efficiencytime has reduced terms through by a similar thedomain amount of HP.as Figure the simulation 13 shows tool that predicts. the climb time has reduced by a similar amount as the simulation tool predicts.

Figure 13. Time to climb vs. hybrid proportion at significantly improved propeller efficiency.

Figure 13. Time to climb vs. hybrid proportion at significantly improved propeller efficiency. Figure 13. Time to climb vs. hybrid proportion at significantly improved propeller efficiency.

Aerospace 2017, 4, 45 14 of 22

Table3 shows the analytic results data from Figure 13, where the propeller efficiency assumed increases from 0.65 to 0.95 along the HP domain, against measured times from X-Plane simulation flights. The times for each completely independent analysis method are now reasonably close, and the trend in changing to the distributed propulsion (three propeller) hybrid layout is very distinct. It is stressed that changes to the X-Plane model between the conventional and hybrid propulsion models created were the empty weight and the extra propellers and nacelles. The total installed power remained constant for both models.

Table 3. Calculated and simulated climb data.

Climb Time Climb Time Installed Empty Payload Aircraft Model Analytic Simulation Power (kW) Weight 1 (kg) Weight 1 (kg) (minutes) (minutes) AUXX 670 2400 800 9 10.6 Conventional AUXX Hybrid 670 3300 800 7.4 8.6 1 Weight given as mass in kilograms including nominal mission fuel.

These results should be interpreted as trends only, as the input parameters are coarse estimates, however the underlying theory and technique for the analysis is based on fundamentals and well known properties. These trends reiterate that hybrid electric aircraft propulsion has many compelling advantages which can improve aircraft performance for particular mission requirements. These include:

• High power/weight (where battery storage capacity requirements are low); • Ease of adding propeller area for any given installed power, allowing greater propulsive efficiency under particular conditions.

While at the same time reducing direct operating costs for energy.

3.3. All-Electric From a survey of conventional aircraft types suitable for the skydiving mission as considered here indicates that a reasonable operating empty weight (the basic weight of an aircraft including the crew, all fluids necessary for operation such as engine oil, engine coolant, water, unusable fuel and all operator items and equipment required for flight but excluding usable fuel and the payload) [17], for this class of aircraft is approximately 2500 kg. Calculating the required energy for the mission and using the lithium battery technology mass specific energy of approximately 200 Watt-hours per kilogram [7] and suitable allowances for other electric power system components yields a take-off weight within the maximum allowable for typical candidate example aircraft types. This calculation takes into account performance variations according to the power usage at any mass according to the installed components. This result concurs with the fact that present state of the art all-electric aircraft such as Pipistrel “Alpha-Electro” are capable of carrying a payload of two people with a short endurance of 1 h in circuit climb and descent mission profile [18]. An all-electric skydiving lift aircraft has been found to be viable using current state of the art EPS technology from an aerodynamic standpoint given the condition that the battery is replaced or recharged for each mission. The primary parameter governing the aerodynamic viability is the weight of the battery required. In this case, given the short overall mission time, a battery mass is feasible which in addition to the payload is within the aircraft load carrying capability. Aerospace 2017, 4, 45 15 of 22

4. X-Plane Modeling

AerospaceX-Plane 2017, 4, is 45 a flight simulation and modeling software tool available online and commercially.15 of 22 In this section X-Plane simulation was conducted to corroborate the findings, of the analytical analysis tooltool detaileddetailed above above and and to contrast to contrast propulsion propulsion configuration configuration examples. examples. Also provided Also isprovided a comparison is a ofcomparison expected performanceof expected performance with a well-known with a well-known legacy type. legacy The softwaretype. The can software provide can high provide accuracy high Aerospacesimulationaccuracy 2017 simulation [, 194, 45] and uses [19] bladeand uses element blade theory element according theory according to specified to geometryspecified geometry to integrate to resultingintegrate15 of 22 forces,resulting accelerations forces, accelerations and consequent and flightconsequent behavior. flight “FAA behavior. Certifications: “FAA X-PlaneCertifications: can provide X-Plane Federal can Aviationtoolprovide detailed Federal Administration above Aviation and (FAA)-certifiedAdministra to contrasttion propulsion simulation(FAA)-certified configuration and vehiclesimulation models. examples. and vehicle This allowsAlso models. provided researchers This allows is toa comparisonachieveresearchers high toof levels expectedachieve of confidence high performance levels in of simulation withconfidence a well-known results” in simulation [20 legacy]. Three type.results X-Plane The.” [20] software models Three were can X-Plane provide developed models high to accuracysimulatewere developed relevantsimulation to configurations simulate [19] and relevant uses of blade aircraft configurations element for comparison. theory of aircraft according for comparison.to specified geometry to integrate resulting forces, accelerations and consequent flight behavior. “FAA Certifications: X-Plane can 1.1. AUXXAUXX Hybrid Electric 1× 1× 224224 kW kW Turbo-Prop Turbo-Prop plus plus 2× 2× 224224 kW kW Electric Electric Motors, Motors, Figures Figures 14–16; 14–16 ; provide Federal Aviation Administration (FAA)-certified simulation and vehicle models. This allows 2. AUXXAUXX Conventional Single Turb Turbo-propo-prop 670 kW, FigureFigure 1717;; researchers to achieve high levels of confidence in simulation results.” [20] Three X-Plane models 3. Cessna208Cessna208 Grand Caravan equipped with 900 HP (670 kW) engine, Figure 1818.. were developed to simulate relevant configurations of aircraft for comparison. Figure 14 is a screenshot from X-Plane “Planemaker”“Planemaker” software environment showing the AUXX 1. AUXX Hybrid Electric 1× 224 kW Turbo-Prop plus 2× 224 kW Electric Motors, Figures 14–16; hybrid aircraft model.model.

2. AUXXThe Planemaker Conventional model Single is flown Turb ino-prop the X-Plane 670 kW, simulation Figure 17; environment in the same way as any

3.ordinaryCessna208 computer Grand flight Caravan simulator. equipped Aerodynamics with 900 HP are (670 modeled kW) engine, using Figure the physical 18. features of the PlanemakerFigure 14 mode,l is a screenshot including from effects X-Plane of wing “Pla sections,nemaker” planforms, software weight environment and balance showing etc. Thethe AUXX hybridHybrid aircraft is shown model. in Figure 15.

Figure 14. X-Plane “Plane-Maker” AUXX hybrid model.

The Planemaker model is flown in the X-Plane simulation environment in the same way as any ordinary computer flight simulator. Aerodynamics are modeled using the physical features of the Planemaker mode,l includingFigure effect 14. X-Planes of wing “Plane-Maker” sections, planforms, AUXX hybrid weight model. and balance etc. The AUXX Hybrid is shown in FigureFigure 15. 14. X-Plane “Plane-Maker” AUXX hybrid model. The Planemaker model is flown in the X-Plane simulation environment in the same way as any ordinary computer flight simulator. Aerodynamics are modeled using the physical features of the Planemaker mode,l including effects of wing sections, planforms, weight and balance etc. The AUXX Hybrid is shown in Figure 15.

Figure 15. X-Plane AUXX Hybrid unrendered model in simulation flight. Figure 15. X-Plane AUXX Hybrid unrendered model in simulation flight. X-Plane can generate well-rendered visual models by draping 2-D image files over a 3-D Planemaker-created aircraft, or by importing 3D graphic files. The aircraft model details rendered in this way do not affectFigure the 15. flightX-Plane physics AUXX Hybridbut are unrendered a useful cosmetic model in enhancement simulation flight. depicted in Figures 16 and 17. X-Plane can generate well-rendered visual models by draping 2-D image files over a 3-D Planemaker-created aircraft, or by importing 3D graphic files. The aircraft model details rendered in this way do not affect the flight physics but are a useful cosmetic enhancement depicted in Figures 16 and 17.

Aerospace 2017, 4, 45 16 of 22 Aerospace 2017, 4, 45 16 of 22 AerospaceAerospace 20172017,, 44,, 4545 1616 ofof 2222

Figure 16. AUXX hybrid rendered model for X-Plane simulation. Figure 16. AUXX hybrid rendered model for X-Plane simulation. FigureFigure 16.16. AUXXAUXX hybridhybrid renderedrendered modelmodel forfor X-PlaneX-Plane simulation.simulation.

Figure 17. AUXX conventional rendered model for X-Plane simulation. Figure 17. AUXX conventional rendered model for X-Plane simulation. FigureFigure 17. 17.AUXX AUXX conventionalconventional rendered model for for X-Plane X-Plane simulation. simulation. In order to compare the performance of the AUXX models in X-Plane with an existing type, a CessnaInIn order208order Grand toto comparecompare Caravan thethe model performanceperformance was employed. ofof thethe AU AUThisXXXX model modelsmodels was inin modified X-PlaneX-Plane withwithto incorporate anan existingexisting the type,type, same aa CessnapropulsionCessnaX-Plane 208208 GrandGrandlayouts can generate CaravanCaravan as the well-renderedAUXX modelmodel wasbutwas otherwise employed.employed. visual re modelsThisThismained modelmodel by standard. was drapingwas modifiedmodified The 2-D C208 to imageto incorporateincorporate model files is over shown thethe samesame a 3-Din Planemaker-createdpropulsionFigurepropulsion 18. layoutslayouts aircraft, asas thethe orAUXXAUXX by importing butbut otherwiseotherwise 3D graphic reremainedmained files. standard. Thestandard. aircraft TheThe model C208C208 details modelmodel rendered isis shownshown in thisinin wayFigureFigure do not 18.18. affect the flight physics but are a useful cosmetic enhancement depicted in Figures 16 and 17.

Figure 18. X-plane C208 model in simulated flight. FigureFigure 18.18. X-planeX-plane C208C208 modelmodel inin simulatedsimulated flight.flight. Each model was flownFigure in the 18.simulatorX-plane from C208 take-o modelff in at simulated maximum flight. rate of climb to 4267 m (14,000 ft) altitude.EachEach modelmodel The waswasaverage flownflown climb inin thethe rate simulatorsimulator was observed fromfrom take-otake-o forffff ateachat maximummaximum aircraft rateratemodel. ofof climbclimb The totoflight 42674267 simulatormm (14,000(14,000 ft)ft) altitude.altitude. TheThe averageaverage climbclimb raterate waswas observedobserved forfor eacheach aircraftaircraft model.model. TheThe flightflight simulatorsimulator

Aerospace 2017, 4, 45 17 of 22

In order to compare the performance of the AUXX models in X-Plane with an existing type, a Cessna 208 Grand Caravan model was employed. This model was modified to incorporate the same propulsion layouts as the AUXX but otherwise remained standard. The C208 model is shown in Figure 18. AerospaceEach 2017 model, 4, 45 was flown in the simulator from take-off at maximum rate of climb to 426717 of m 22 (14,000 ft) altitude. The average climb rate was observed for each aircraft model. The flight simulator software includes a full “autopilot” facility which enables straight-forward and repeatable simulated software includes a full “autopilot” facility which enables straight-forward and repeatable simulated flights. In this work, the autopilot setting maintained a given pitch and roll angle throughout the flights. In this work, the autopilot setting maintained a given pitch and roll angle throughout the climb. climb. The roll angle was set as close as possible to zero, while the pitch angle setting was decided The roll angle was set as close as possible to zero, while the pitch angle setting was decided through through experiment to give best climb speed. It was noted that the indicated airspeed reduced experiment to give best climb speed. It was noted that the indicated airspeed reduced slightly during slightly during the climb for all models, while the rate of climb steadily reduced as expected. the climb for all models, while the rate of climb steadily reduced as expected. Of particular interest in determining relative performance and efficiency is the effect of propeller Of particular interest in determining relative performance and efficiency is the effect of propeller slipstream “scrubbing”. “The part of the body, nacelle, wing or other airplane component which is slipstream “scrubbing”. “The part of the body, nacelle, wing or other airplane component which is directly located in the slipstream will experience a change in drag, called scrubbing drag. This effect directly located in the slipstream will experience a change in drag, called scrubbing drag. This effect can be “counted” as a change (increase) in airplane drag or as a change (decrease) in the installed can be “counted” as a change (increase) in airplane drag or as a change (decrease) in the installed thrust of the propeller.” [21]. thrust of the propeller” [21]. The AUXX design requirements result in a body with very large fuselage frontal area ratio of the The AUXX design requirements result in a body with very large fuselage frontal area ratio of propeller disk area compared to most similarly configured types. Although the fuselage is very well the propeller disk area compared to most similarly configured types. Although the fuselage is very contoured, it inevitably has larger pressure gradients and wetted area contributing to drag, especially well contoured, it inevitably has larger pressure gradients and wetted area contributing to drag, scrubbing drag, than comparative size and weight aircraft in the class which have much slimmer especially scrubbing drag, than comparative size and weight aircraft in the class which have much fuselages. Figure 19 shows the relative frontal cross-section area of the AUXX compared to legacy slimmer fuselages. Figure 19 shows the relative frontal cross-section area of the AUXX compared to types. legacy types.

FigureFigure 19.19. AUXXAUXX frontalfrontal areaarea comparison.comparison.

TheThe advantagesadvantages ofof thethe conventionalconventional singlesingle frontfront mountedmounted engineengine configurationconfiguration areare clearclear andand acknowledgedacknowledged inin thethe designdesign requirements.requirements. CommercialCommercial experienceexperience inin thethe missionmission confirmsconfirms thatthat thethe operatingoperating andand capitalcapital costscosts ofof twintwin engineengine (ICE)(ICE) configurationsconfigurations mustmust bebe avoided.avoided. ItIt isis alsoalso aa knownknown problemproblem ofof lightlight twintwin turbineturbine poweredpowered aircraftaircraft thatthat typicaltypical locationlocation ofof wingwing mountedmounted enginesengines placesplaces significantsignificant risksrisks ofof uncontaineduncontained turbineturbine bladeblade failurefailure affecting affecting the the cabin cabin [ 22[22].]. However,However, the the ability ability to locateto locate electric electric motors motors and propellerand propeller units onunits the wingson the presents wings anpresents excellent an opportunityexcellent opportunity to improve to the improve propulsive the propulsive efficiency, ef byficiency, reducing by propeller reducing disk propeller loading disk and loading scrubbing and drag,scrubbing while drag, retaining while the retaining advantage the advantage of single front of single mounted front enginemounted placement. engine placement. The electric The motor electric is lowmotor temperature, is low temperature, low speed low and speed low cost and by low comparison cost by comparison to equivalent to powerequivalent gas turbinepower equipment,gas turbine therebyequipment, eliminating thereby the eliminating gas turbine the blade gas failureturbine risk blade while failure providing risk while an economical, providing higher-efficiency an economical, distributedhigher-efficiency propulsion distributed installation. propulsion installation. AnotherAnother advantageadvantage toto bebe establishedestablished inin thethe simulationsimulation resultsresults isis thethe performanceperformance ofof thethe electricelectric motorsmotors atat altitude. altitude. The The X-Plane X-Plane simulation simulation uses uses a gas a turbinegas turbine flat-rated flat-rated to 10,000 to ft.10,000 Above ft. thisAbove altitude, this altitude, the maximum power available reduces in proportion to reducing air density. The electric motor power output remains constant with altitude. The results for average climb rate of the simulated aircraft and key parameters are shown in Table 4.

Aerospace 2017, 4, 45 18 of 22 the maximum power available reduces in proportion to reducing air density. The electric motor power output remains constant with altitude. The results for average climb rate of the simulated aircraft and key parameters are shown in Table4.

Table 4. X-Plane aircraft simulation model parameters and climb rate.

X-Plane Aircraft Installed Power Empty Weight 1 Payload Weight 1 Average Climb Rate Model (kW) (kg) (kg) to 14,000 ft (ft/min) AUXX 670 2400 800 1320 Conventional AUXX Hybrid 2 670 3300 800 1620 C208 Conventional 670 2200 800 1550 C208 Hybrid 2 670 3000 800 2150 Real Aircraft Data 850HP C208 3 635 2200 1800 * 1460 4 1 2 Weight given as mass in kilograms including nominal mission fuel; HP = 0.66, 1× Front mount gas turbine plus 2× wing-mounted electric including battery and electrical components; 3 [23,24]; 4 * Initial rate at maximum gross weight.

These results show a clear trend of improved climb rate with the application of HEP at an HP of 0.67 even with the necessary increase in weight. Since only the weight and the number and location of propellers was changed in the model, it has to be assumed the improved performance in simulation was due to increased efficiency.

4.1. Multi-Role The analysis above was concentrated on the particular mission of high frequency climb and descent cycles with an assumption that the electrical propulsion battery would be changed out or recharged for each cycle. The feasibility of the mission is shown as viable with an attendant reduction in fuel consumption and, therefore, direct operating cost and emissions. However, varying the mission requirements for the in-service aircraft, perhaps for regional passenger transport or freight operation, will render the range and endurance inadequate when using any significant degree of hybridisation due to the limitation of current and near-term battery energy density technology. Such a reduction in role versatility is likely to prevent an operator investing in this new technology aircraft equipped with hybrid electric components. The benefits of the hybrid electric installation remain clear; the redundancy of including two separate propulsion systems, but without the problems inherent in conventional twin engine layouts; the possibility of utilizing non-hydrocarbon energy sources; the improved propulsion and aerodynamic efficiency etc. So, how can these conflicting objectives be resolved in order to allow the best future concepts to emerge?

4.2. Modularity A solution to the problem is to introduce a new paradigm for aircraft propulsion modularity. Different roles and missions can be achieved using mixtures of basic component parts all designed to fit standardized interfaces. The concept is currently used for example on military aircraft external stores. A standard rack can carry a large variety of payloads, munitions, sensor pods, fuel tanks etc. For new propulsion, where a range of different hybridisation degrees, sizes of engine, battery packs etc., allow a variety of different performance outcomes, this concept can allow economic viability. As an example, using the aircraft modeled above with a 223 kW ICE installed, and 447 kW of EM, we note that the effective range is severely limited. The battery size may be reduced to allow carriage of more fuel, but the remaining propulsive energy of the EM is then reduced, and the propulsive power of the ICE is inadequate for safe flight across all phases. In this case, removing the battery and installing a larger ICE to support electrical energy demands of the EMs would provide acceptable Aerospace 2017, 4, 45 19 of 22 power and range, though at potentially lower propulsive efficiency due to the removal of the wing mounted propellers With some compromise in overall structural mass efficiency, it is a straight-forward engineering process to provide a firewall forward change-out of the ICE components. Such a task could realistically be achieved within minutes if the design were suitable. It may be useful for some missions to remove the EM units or fit different sizes etc. It is the capability in practice for such module interchange in the airworthiness and certification process which would currently cause the most delay. The legacy airworthiness and certification process was not designed for and is unfit for this type of approach, but just as that process was developed according to the requirements and constraints of the time, so a new system can be developed for new requirements and constraints. An operator could invest in a set of components based on a standard fuselage and tail group. Various ICE, EM, and wing group parts suiting different missions could be interchanged as necessary for the best cost/performance outcomes.

5. Future Analysis The aircraft concepts in this paper have been explored according to fundamental mission requirements and the constraints of propulsion technology. Future analysis will extend this approach into a whole of system design domain using advanced conceptual design methodologies to evaluate the performance attributes of these hybrid-electric propulsion concepts. It will approach this using an analysis based on conceptual design methodologies such as the concept design analysis (CODA) [13], quality function deployment (QFD) [25,26] , decision matrices (DM) [27,28] and value driven design (VDD) [29,30] approaches. The methodology will establish mission requirements and needs in relation to provisions for system performance, sustainability, certification, weight and cost metrics. Given that the system solution space may include hybrid combinations of ICE and EP configured in multiple multi-motor architectures, and/or operated to various battery charging strategies, it is likely that numerous potential concepts will result. One particular method used has been the application of general morphological analysis techniques [31,32] to evaluate various technology combinations and options. This is achieved using a concept morphological matrix approach which develops technology vectors which can be evaluated within integration requirements. Different concepts are therefore developed by combining various possible solutions exhaustively to ensure that a rigorous examination of the design space is conducted. Later steps in this methodology will therefore involve selection of the “best candidate” concepts, and optimization of the solution based on change propagation analysis [33–36]. This change propagation analysis method will also use matrix methods combined with various visualization techniques to assess the impact of changes resulting from HEP related aircraft modifications. This analysis methodology will assess the impact of change propagation on integration risk and evaluates technology readiness level. In addition, this methodology can assist in evaluating technical performance attributes, sustainability and value proposition of the HEP solution.

Safety and Certification Commercially operated aircraft are subject to stringent certification and standards control. In many jurisdictions, skydiving lift aircraft comply to FAA Part 23 normal category standards. This category allows up to 5700 kg maximum take-off weight (MTOW) [28]. Aviation certification and standards and regulations often lag technological development and may become anachronous in view of disruptive technologies. Yet, the foundation of these processes remains vital to the ongoing safety of aviation and will not disappear. However, the adoption of conceptual design methodologies as described above provides a means to assess those certification requirements impacted by HEP aircraft modifications. This methodology would build on change propagation analysis techniques as described above. It is anticipated that this would a research project in its own right which would inform regulatory changes accounting for these new propulsion technologies and maintaining the salient hard-won safety principles developed over the past 100 years. Aerospace 2017, 4, 45 20 of 22

6. Conclusions A novel aircraft design has been analysed as a case study to highlight the positive advantages of hybrid electric propulsion for aircraft. Whereas negative compromises of electric propulsion remain significant due to increased system weight compared to pure internal combustion alternatives, various efficiencies and advantages can be shown to allow overall performance improvement for particular missions and some new aircraft roles become viable only as a result of hybrid electric propulsion. It has been shown how a short duration high power mission such as skydiving can use HEP equipped aircraft to reduce fuel consumption, and how even simply-configured distributed propulsion benefits the efficiency to offset the problematic aspects of increased HEP system weight. The use of a hybrid electric propulsion system could yield a viable bridge between legacy certification standards and new evolving standards for fully electric aircraft. The way in which the traditional ICE is integrated with electrical components can provide enhanced performance as well as system redundancy. This changes the whole risk profile for the aircraft capital investment, development and operation. It is evident that more and more hybrid and electric aircraft concepts are emerging from small light sport types through to intercontinental heavy transport. It is inevitable that the new opportunities for aircraft design will result in a range of new types with technical and commercial viability, but requiring reform and advancement of regulatory and certification systems. Many studies of electric and distributed propulsion for aircraft look at the opportunities which may only exist once significant further battery storage technology and other future EPS improvements take place. The study presented here exemplifies the fact that HEP technology can have utility in aviation today with current levels of technology. Moreover, the study shows that new aircraft developed using current HEP technology can lead progress, and will only benefit as future technology emerges, rather than being superseded.

Acknowledgments: The subjects of these case studies are private development projects used by courtesy of the developers. The “AUXX” Skydiving aircraft project is owned by Mr Tibor Glesk, Air Ute Pty Ltd., Caloundra, Queensland, Australia. Author Contributions: Richard Glassock: Richard is a Research Fellow at the University of Nottingham Institute for Aerospace Technology. Richard wrote the body of the manuscript and developed the MATLAB and X-Plane modelling tools, as well as the 3D CAD models used in design development and illustrations for the case study aircraft. Michael Galea: Michael received his PhD in Electrical Machines Design from the University of Nottingham, UK. He is currently a Lecturer in Electrical Machines and Drives within the Power Electronics Machine and Control research group of the University of Nottingham. Michael reviewed the manuscript and provided input for all items regarding electrical systems. Warren Williams: Warren is a professional aeronautical engineer in private practice in Australia and is currently completing a PhD in Aeronautical Engineering. Future Analysis topics of Safety and Certification were contributed by Warren. Tibor Glesk: Tibor is a professional skydiver who has operated commercial skydiving centres around the world for 30 years. Tibor designed the conceptual requirements of the AUXX aircraft, owns and manages the design program. Tibor provided advice and feedback on all areas of commercial and operational requirements for the skydiving aircraft concept. Conflicts of Interest: The aircraft used in this paper is subject to ongoing development projects toward commercialisation. Contribution from private project resources was limited to allowing use of data and figures. No funding, sponsorship or financial consideration of any sort was given.

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