Assessment of Thermo-Electric Power Plants for Rotorcraft Application

This paper assesses a parallel electric hybrid propulsion system utilizing simple and recuperated cycle gas turbine configurations. An adapted engine model capable to reproduce a turboshaft engine steady state and transient operation is built in Simcenter Amesim and used as a baseline for a recuperated engine. The transient operation of the recuperated engine is assessed for different values of heat exchanger effectiveness, quantifying the engine lag and the surge margin reduction which are results of the heat exchanger addition. An oil and gas mission of a twin engine medium helicopter has been used for assessing the parallel hybrid configuration. The thermo-electric system brings a certain level of flexibility allowing for better engine utilization, thus firstly a hybrid configuration based on simple cycle gas turbine scaled down from the baseline engine is assessed in terms of performance and weight. Following the recuperated engine thermo-electric power plant is assessed and the performance enhancement is compared against the simple cycle conventional and hybrid configurations. The results indicate that a recuperated gas turbine based thermo – electric power plant may provide significant fuel economy despite the increased weight. At the same time, the electric power train can be used to compensate for the reduced specific power and potentially for the throttle response change due to the heat exchanger addition.


INTRODUCTION
The aviation industry is recognized as the most rapidly growing source of CO2 emissions given that the 2035 global traffic is fore-casted to be twofold the one in 2016 [1].Rotorcrafts are going to be part of this growth since helicopters are seen as a quick and safe way to transport patients between hospitals [2].Additionally the traffic for passenger transport/air taxi, which has been a marginal activity until now, is expected to boom in the near future with a forecasted two to three-fold increase in the 2015-2020 period [3].Therefore, the rotorcraft fleet contribution to environmental impact is becoming a concern and measures to reduce it are investigated and assessed.The Advisory Council for Aeronautics Research in Europe (ACARE) has established specific environmental goals to be met by the civil aviation industry (including rotorcraft and aircraft) by the year 2020 in order to sustain aviation's present environmental impact after the expected aircraft and rotorcraft fleet growth [4].
The development of new technologies can potentially hinder the environmental emissions of the aviation industry.Several ways to tackle this challenge have been proposed ranging from changing the aircraft operational procedures [4] to system architecture ( [5], [6]).The "More Electric Aircraft" (MEA) concept is expected to be in the forefront of this change, given that already pneumatic and hydraulic aircraft systems are replaced by electric systems [7].On longer term hybrid electric and universally electric aircraft are considered as promising concepts for addressing the NASA N+3 [8] goals and the Strategic Research Innovation Agenda as discussed by Vratny et al. [9].
Thermo-electric configurations bring a certain level of flexibility into the powerplant design and the vehicle as a whole, since the engine may operate in more favourable operating points throughout the mission, providing propulsive power and electric power in several systems at the same time.This flexibility is of interest for rotorcrafts since typical helicopter cruises between 55% to 65% of the maximum continuous power for around 80% to 90% of the helicopter flight [10], meaning that they operate at high specific fuel consumption (sfc).Another interesting aspect is that in the case of hybrid propulsion the OEI (One Engine Inoperative) 1 rating can be achieved by utilizing the battery pack adding to the safety and reducing the need for engine oversizing.
Especially for rotorcrafts, hybrid electric configurations may offer high reliability and lower maintenance cost, mainly due to reducing the gearboxes and transmissions, components that do not have graceful failure modes.These advantages come with increased weight and complexity of the whole propulsion system, thus for assessing the performance of a hybrid propulsion system the increased weight of the helicopter should be considered.
Another way to address the increased specific fuel consumption during part-load operation of helicopters is the recuperated cycle.The deployment of regenerative technology for enhancing the operational capabilities of helicopters, specifically for military use, has been of interest since 1960s [11].Despite the well established benefits in terms of fuel consumption that can accrue from recuperation reliability concerns and the insignificant operating cost savings in an era of low-fuel cost; made this technology less attractive since there was still room for significant improvement of the simplecycle [12].Through the continuous advancement of gas turbines the simple cycle configuration approached its performance limits and new aircraft designs integrating heat exchangers has been proposed (e.g.[13]).This development reignited interest in recuperated turboshaft engines as well.Fakhre et al. [14] demonstrated that the conceptual recuperated helicopter has the potential to significantly improve the maximum attainable range capability, achieving a significant reduction in mission fuel burn but may cause an increase in NOx emissions.
The advantages of the recuperated cycle come with increased weight, complexity and lower specific power due to the heat exchanger pressure losses.The latter is important if an existing engine is considered for modification, since its power ratings are expected to significantly decrease, thus it may not be capable to power existing helicopters, as discussed by Roumeliotis et al. [15].Additionally the dynamic behaviour of the engine is expected to change due to the thermal inertia of the heat exchanger changing the throttle response of the system [16].Despite the fact that recuperated configurations have been proposed for rotorcraft propulsion, recuperated cycles as part of a thermo-electric power plant have not be assessed in the literature.A thermo-electric powerplant utilizing a recuperated cycle can address the negative effect of heat exchanger pressure losses to the engine power output, since the electrical power train can provide the power boost needed at specific mission segment, making this concept more interesting.In this case existing engines can be modified decreasing the development time and cost.Additionally a hybrid -electric configuration can compensate for the thermal inertia effect during acceleration by utilizing the electric power system since electric motors can produce their torque within a few milliseconds [17].The aim of this work is to give a deeper insight in the performance and operability characteristics of thermo-electric power plant based on recuperated gas turbine for helicopter applications.A parallel hybrid configuration is selected since no distributed propulsion is considered, thus eliminating the need for adding the weight of a generator ( [15], [18]).
For assessing the hybrid configurations for rotorcraft propulsion relevant dynamic models are built in Simcenter Amesim.A simple cycle engine adapted on publicly available experimental data of a T700-GE-700 is used as the baseline.Following a recuperated version of the engine is developed and its performance is assessed for different values of heat exchanger effectiveness.The transient operation of the recuperated engine is also simulated for assessing its operability and reaction time.
Having established a recuperated version of the engine an integrated parallel hybrid configuration utilizing both the simple cycle and the recuperated configuration models is built in Simcenter Amesim, allowing for the dynamic simulation and performance assessment of both the turboshaft engine and the electric power train at the same time.

SIMCENTER AMESIM GAS TURBINE LIBRARY
The Simcenter Amesim Gas Turbine library consists of components such as compressor, turbine, combustor, ducts etc.The modular nature of a gas turbine engine makes the identification of components and connections a straight forward procedure.
In addition to the usual components, volume components can be used for taking into account the volume dynamics effect on engine transient operation and variable orifices to simulate bleed off valves (BOV).Components from other libraries such as shafts for accounting for shaft dynamics, propellers, gearboxes, generators, heat exchangers, pumps can be combined with the gas turbine library components to allow the modelling of any gas turbine configuration and of integrated systems as well.
The compressor and turbine modelling is based on the well established rigorous entropy calculations described by Walsh and Fletcher [20] and Kurzke [21].The variable geometry effect on turbomachinery map is simulated by applying the methodology described by Kurzke [21].Compressor and turbine maps correction for accounting working medium composition (e.g.humidity for compressor and fuel synthesis for turbine) is done utilizing the method suggested in AGARD AR-332 [22].This feature can be activated / deactivated.For simulating turbomachinery components degradation the health indices concept as introduced by Stamatis et al. [23] has been implemented.
Suitable health indices can be used for simulating specific components faults [24].
The volume dynamics and the heat transfer between the gas and the surrounding parts of the engine are represented by the chamber component with heat exchanges.The heat flow can be modelled utilizing the Simcenter Amesim thermal library components and the relevant materials library.The ducts are modelled as orifices of specific flow coefficient, where the flow coefficient can be calculated via available geometrical data or imposed by the user.This approach allows for varying pressure losses throughout operation, as a function of corrected mass flow.The combustion is modelled utilizing chemical kinetics and the relevant reaction rate coefficient [25].Any fuel composition can be applied and the combustion products are calculated based on the stoichiometric equation.The combustion chamber components available are: the primary zone accounting for the formation of the flame kernel, combustion and pressure drop and the secondary zone accounting for further chemical reactions and pressure drop.Both components can account for heat transfer applying thermal library components, while cooling holes can be considered too.The secondary zone can be further discretized utilizing several secondary zone components for representing the dilution zone.The fluid properties are calculated according to the NASA CEA data [26].
The ambient air synthesis can be defined by the user in molar or mass basis.The ambient air conditions can be defined by the user or calculated via the atmosphere component.International standard atmosphere (ISA) [27], U.S. standard atmosphere selection.The ISA non-standard days are also available (hot, cold, tropical and polar) for assessing the system operation in the extremes of the design envelope.
By using the gas turbine library any engine model can be built.The modular nature of the gas turbine engine makes this a straight forward process.In order to relate flow quantities at the inlet and outlet of individual engine components, the equations of mass and energy conservation along with species concentration conservation are employed, while shaft components impose torque calculations.The equations are solved through a well established integrator switching automatically between several algorithms depending on the numerical stiffness of the system [31], [32], [33].
The gas turbine library set of components can be combined with a plethora of components of other libraries such as electrical, control and thermal, allowing the modelling of conventional and unconventional engines, the analysis of integrated systems and the application of different control strategies.For streamlining the engine model building process a compressor and turbine map scaling tool is available.The engine design point can be derived from the open literature [34], measurement data [35] or design calculations based on technology level [36].The map scaling tool (Figure 1) allows for the point and click selection of the initial map design point, hence the map scaling factors [37] can be directly calculated.The simple cycle engine is based on the T700-GE-700, a turboshaft engine broadly used in Twin Engine Medium (TEM) helicopters.A model adapted to available experimental measurements [19] is built and assessed.Then the simple cycle engine is acting as the baseline for a recuperated engine.

Simple Cycle Engine
The T700-GE-700 is the first production model of the T700 family (first engine delivered in early 1978) [38].It consists of a five-stage axial and a single-stage centrifugal flow compressor; a low-fuel-pressure, annular combustion chamber; a twostage axial flow gas generator turbine; and a two-stage uncooled independent power turbine [38].The engine model built can be seen in Figure 2, utilizing fuel flow as the setting parameter and applying the maximum allowable turbine inlet temperature (T41), as defined at stator exit, as limiter.the CT are extracted from the HPC exit, assuming a bleed percentage of 10% relative to the inlet flow, following the cooling flow data presented by Ballin [19].The cooling flow split is selected following the equivalent single stage cooled turbine logic [20], [39].The work potential of the cooling flow is defined equal to 0.6 based on the value suggested by Kurzke [39] for two stage cooled turbines.
For ensuring low power stable operation the first and second stator vanes of the axial compressor are of variable geometry (VGVs) and starting air is bled off from the axial compressor exit through a bleed off valve (BOV).Both VGVs and BOV position is a function of ambient temperature and rotational speed and they are activated when the corrected gas generator speed is less than 87% of the nominal one.The VGVs are modelled utilizing the relevant mass flow correction as a function of angle [21], assuming no penalty on efficiency and pressure ratio according to the measurement data discussed in [35].The BOV is modelled through a variable area orifice component.
Kerosene (C10H20) is used as fuel in this study.
The model is adapted to the experimental data presented by Ballin [19] and depicted in Table 1.The operating point for scaling the model and maps is selected the one with the maximum torque.Following a model adaptation to the engine available data utilizing components modification factors [35] was performed.As seen in Figure 3 the adapted engine model reproduces the engine steady state experimental data with very good accuracy throughout the whole operating envelope.The model deviates from the experimental data when operation close to idle is considered.This is expected since there are no public domain data available for the engine component characteristics on the idle operating regime.The turbomachinery component low rotation speed isolines, VGVs and BOV schedules, to name some parameters, are of high importance when operation close to idle is considered.In this case, generic component maps and schedules are used.For this size of engine the gas dynamics effects are expected to be small [41].The heat soakage effect is expected to be small according to the results presented by Ballin [19] and it is not considered in this model.The heat soakage effect can be easily applied utilizing components from the thermal and material libraries.
The transient behaviour of the model for the case of a fuel step increase can be seen in Figure 4.The gas generator speed, compressor discharge pressure and torque variation are presented along with calculated data presented by Ballin [19].

Recuperated Engine
In this study the test case engine is acting as the baseline for the recuperated engine.Recuperated engines have been considered for rotorcraft configurations (e.g.[12], [14]) and assessed from a performance point of view.The simple cycle engine model built is easily changed to a recuperated cycle utilizing the relevant available heat exchanger components.Simcenter Amesim offers several levels of detail for modelling the heat exchanger.The simplest model is selected, applying off design change of effectiveness based on eq. ( 1) [20]: Basic sizing data is needed for the heat exchanger for modelling its transient operation and for evaluating its weight, namely the characteristic length, area, volume and mass.These values are calculated according to the simplified method described by Juhasz [42] and adopting the values suggested by the author, specifically: heat transfer coefficient equal to 142 W/(m2 K), density equal to 3900kg/m 3 and packing density of 3281m 2 /m 3 .The heat exchanger pressure drop at design point is assumed 4% for the cold side and 5% for the hot side including the exit duct losses, in line with the values suggested in [20].The effectiveness of the recuperator is expected to be in the range of 0.4 to 0.74 according to previous designs presented by McDonald et al [43].
A parametric analysis assuming design point effectiveness equal to 0.4, 0.6 and 0.8 is performed.The maximum continuous rating power (MCR) and max intermittent rating power 2 of the recuperated engines are defined in accordance with the simple cycle engine T41 values.The simple cycle MCR and max intermittent rating are defined in accordance to the values presented in [44].Having established the simple cycle power ratings the relevant T41values are established as presented in Figure 5.As seen in Figure 5 the imposed pressure losses by the recuperator have a detrimental effect on engine power.The overall performance for the two power ratings and for the recuperated and simple cycle engines is presented in Table 2.
As seen the recuperated configurations have a max intermittent rating power decreased by approximately 200kW compared to the simple cycle.The MCR power is decreased by more than 120kW compared to the simple cycle engine.This means that these recuperated engines cannot be used on the same rotorcraft as the simple cycle engine.Hybrid propulsion systems may provide the additional power boost, negating this disadvantage.As expected, recuperation produces a significant betterment in terms of specific fuel consumption.The improvement is higher when the engine operates at part load as seen in Figure 6.The decrease of the recuperated engines sfc compared to the simple cycle sfc is depicted in Figure 7. Values of recuperation effectiveness in the range of 40-60% are considered plausible and in these cases a mean sfc reduction ranging from 10% to 15% can be expected.The dynamic behaviour of recuperated engines for rotorcraft propulsion has not been assessed despite the fact that the recuperator is expected to affect the operability and response of the engine [16].As seen in Figure 8, the addition of the heat exchanger changes the response time of the gas turbine to external disturbances, as expected, due to the recuperator volume and thermal inertia.Higher effectiveness means larger volumes and masses resulting to longer engine reaction times.The change on the engine throttle behaviour may be addressed by suitable control system, since suitable over-fueling should be addressed or in the case of a hybrid system by utilizing the electric power train to provide the power reducing the lag.
This behaviour should be assessed when recuperated engines are considered, since there are specific manoeuvres that may be affected by an engine power lag.This lag is a strong function of thermal inertia, volumes and control system; hence for fully assessing recuperator effect on the overall engine transient operation more detailed heat exchanger design is needed along with a suitably design control system.
Nevertheless, the recuperated model demonstrated a physically consistent behaviour and can be used for deriving general trends.Recuperation increases the pressure ratio of the engine for a specific operating point due to the increased pressure losses; hence it is negatively affecting both compressors surge margin.This negative effect becomes more profound when transient operation is considered.The acceleration-deceleration manoeuvre depicted in Figure 9 has a significant negative effect on surge margin for both engine compressors as depicted in Figure 10 and Figure 11.The effect of the heat exchanger thermal inertia on compressors operating line is significantly more profound than the effect of shaft inertia, moving the operating line towards surge and highlighting the importance of examining the engine operability along with the performance when new configurations or modifications that include heat exchangers are assessed.For assessing a hybrid propulsion configuration in terms of performance and fuel consumption the power train weight should be evaluated.

Gas Turbine Weight Estimation
The gas turbine engine may be sized down, given that the batteries can provide boosting power, or sized up if a single gas turbine configuration is selected.For accounting the engine scaling effect on the engine weight ATLAS, a Cranfield University in-house software [45] for engine sizing and preliminary structural assessment is used.
ATLAS utilizes a component based approach in order to derive the overall "bare-engine" weight.Based on gas turbine theory fundamentals, it performs gas path designation and preliminary design for each of the main engine components.
Design considerations such as stage loading and de Haller number limitations, blockage factors, reaction targets and flow Mach numbers are included in order to determine the required technology level.Each engine component is accompanied by the rest of the parts necessary to complete a realistic engine module.The component designs combined with the appropriate gas turbine materials library are used for the engine weight estimation.
For the rotating discs, an approach similar to the one used in NASA WATE method is applied [46].In this approach, the target is to find the minimum volume, and subsequently the minimum weight disk that satisfies the stress limitations imposed by its shape and its material.The main assumption here is that the disks are consisted by concentric rings while the calculation is performed based on the Sums & Differences method [45].The structural evaluation follows the Goodman approach.
The 2D engine cutaway generated by ATLAS for the test case engine utilizing the aerothermal data calculated by the model for the maximum intermittent rating is depicted in Figure 12 along with the original engine layout.

Figure 12: Engine cutaway comparison
The geometry obtained with ATLAS is comparable to the actual engine layout.
The overall length and weight predicted by ATLAS for the test case engine are in good agreement with the actual length and weight reported in [44] (Table 3).Given that ATLAS can predict the engine weight utilizing performance data with acceptable accuracy; it is used herein to assess the engine weight change when engine scaling down is considered.

Electric Power Train Weight Estimation
For evaluating the electric power train, the weight of each component should be calculated.The electric components weight is a strong function of the technology level.
For the batteries, current state-of-the art Li-ion batteries have specific energy in the range of 200 Wh/kg, while future developments are expected to increase this value to 2000 Wh/kg [47].In order to estimate batteries weight, the overall amount of energy (Etot) is used, along with the energy density (ρbatt).Additionally a service factor kadd equal to 1.3 is introduced, in order to take into account mass coming from harnesses and casing [48].
The inverter aims at changing direct current into alternating current, or viceversa, depending on the application.Its mass can be related to the amount of power it has to handle, and can be computed knowing the power density (ρinv).According to [49] a cable mass factor kcables equal to 1.2 is added to account for wires and electrical connections.
Inverters usually suffer from low power density, but those based on SiC material can reach 7 kW/kg [51].The power density of the inverter is in the range of 2 kW/kg, while future values are expected to be as high as 16 kW/kg [50].Concerning the motors and generators the method adopted to compute the mass and the overall geometry is taken from [49] utilizing data from [52].The design power along with the design and maximum rotational speed are inputs.The number of phases and poles, along with material components density are considered design variables.This input is used for calculating the electric machine components (stator, rotor, shaft and armature) sizes and weights.The overall mass is obtained summing the components weights.A service mass equal to 13% of the total is added to take into account the added weight of cables and mountings.

THERMO-ELECTRIC POWERPLANT ASSESSMENT Test Case Mission
For assessing the thermo-electric powerplant a specific Twin Engine Medium helicopter and a representative Oil and Gas (OAG) mission are considered.HECTOR has been used for deriving the power demand for the specified mission.HECTOR is an inhouse rotorcraft comprehensive code [53] that has been used to reproduce with good accuracy power requirements for TEM helicopters [54], [55].
The OAG mission schedule assumes that the helicopter takes off from De Kooy Airfield in Den Helder, the Netherlands while carrying a specified payload.The helicopter subsequently travels towards a designated offshore oil/gas platform (oil rig 1) where it lands and drops off the on-board payload.The helicopter then travels towards a second offshore oil/gas platform (oil rig 2) where it picks up another useful payload A parallel hybrid propulsion system for helicopters is examined herein with the partial electric drive of the main rotor supplementing the engine power.It is assumed that the twin-engine configuration is maintained, ensuring redundancy.
The propulsion system consists of two power sources, each consisting of a gas turbine mechanically connected to the motor through a gearbox that acts as a power splitter, while a battery provides power to the motor.The electric motor is used to assist the turboshaft engine when the power request is very high and for charging the battery when surplus of mechanical power is available and the battery stage of charge (SOC) is less than 30%.A relevant model is built in Simcenter Amesim using the Aircraft Electrics and Electric Motor and Drives libraries (Figure 14).
The battery, controller and motor design is done utilizing the Simcenter Amesim tools providing the voltage, energy and power demand for the battery and the power, the design and the maximum rotational speed for the motor and controller.The inverter efficiency is assumed constant and equal to 0.96 [50].The motor design point efficiency is assumed equal to 0.94 [56], and changes at off-design operation.The torque demand is provided by the HECTOR.The gas turbine power throughout the mission is limited by the MCR turbine entry temperature (T41) as defined for the simple cycle.When the T41 becomes equal to the MCR T41, the fuel is stabilized and the battery is supporting the engine, providing the additional power, if needed.
Two configurations are assessed herein, one utilizing simple cycle engines as prime movers and one utilizing the recuperated engines:  The OAG mission with a TOW of 5760 kg is defined as the test case mission, except if specifically stated otherwise.For each configuration, performance parameters such as the gas turbine overall mission efficiency, eq. ( 4), propulsion system overall efficiency, eq. ( 5), energy ratio, eq. ( 6) and the ratio of maximum electric power to propulsive power occurred during the flight, eq. ( 7) are calculated.Additionally a corrected system efficiency is introduced (eq.( 8)) for taking into account that the electric energy production and its transfer to charging stations for charging the propulsion system batteries is not a 100% efficient process.The mission results (electric energy, maximum electric power and fuel consumption) are used for calculating the hybrid propulsion system overall weight.The effect of engine scaling to the engine weight and the weight of the heat exchanger as a function of effectiveness are also considered.

Simple Cycle Configurations Assessment
The energy split for the simple cycle engine configurations versus the engines design MCR is depicted in Figure 15.For this mission the engine has to be sized down to 80% of the baseline engine for the electric support to become necessary.This doesn't mean that an engine sized down to 90% of the baseline engine power will not need a power source for ensuring maximum intermittent and OEI capabilities.In the case of scaled down engines even if no electric power is used during the mission a hybrid system is still needed for the propulsion system to be capable to provide the maximum power ratings.As seen in Figure 16 the hybridization in terms of energy becomes significant when the engine is scaled down to the 70% of the baseline engine power.At the same time the maximum electric power to propulsive power ratio occurred during the mission is 9.5% for the engine scaled down to 80% of baseline power and it increases up to 36% for the engine scaled to 60% of the baseline engine.The gain due to operating point movement is expected to decrease as the engine is becoming smaller.This is due to the performance curve slope decreasing for higher power settings.This is evident when the gas turbine overall mission efficiency is examined.The gas turbine overall efficiency is saturated when the engine is scaled down to approximately 70% of the baseline engine power, as seen in Figure 18.The overall efficiency is increasing given that the electric power is considered as input to the system with a mean efficiency of about 89% (motor and inverter efficiency).
In reality, considering that the electric power may be for example produced from a Combined Cycle power plant, the efficiency conversion from fuel heat energy to electric energy would be in the range of 50 to 60%.If this correction is employed the overall system efficiency gain due to hybridization is reduced considerably, as seen in Figure 18.
This decrease highlights also the fact that utilizing in-flight charging will further compromise the overall hybrid system efficiency, since on top of the Brayton cycle inefficiencies the electric power train inefficiencies would be applied twice.On the other hand in-flight charging may be necessary for reducing batteries weight and for increasing safety.
Having established the electric energy needed the electric power and the engine power the overall propulsion system weight can be calculated utilizing the methods described.The batteries power density is assumed 200Wh/kg and the controller power density 7kW/kg.It should be highlighted that the electric motor and controller are sized so the propulsion system can provide the baseline engine high power ratings.The battery for the engine scaled to 90% of baseline engine power is sized as to ensure maximum power availability for 2.5 min.The batteries mission energy demand is increased by a safety factor of 1.2.The relevant data is presented in Table 4.As seen the weight increases significant as the hybridization increases, making the configurations utilizing gas turbines scaled down to 70% and 60% of the baseline engine power very heavy.

Recuperated Cycle Configurations Assessment
For the recuperated cycle configurations the hybridization is rather small, as seen in Figure 19.The hybridization increases with recuperator effectiveness due to the MCR decreasing.Given that the electric power is providing power at short intervals throughout the mission the dominant efficiency is the gas turbine efficiency, as seen in Figure 20.
The gas turbine efficiency is rather high, despite the fact that the engines are not scaled down.The efficiency increase comes with the penalty of the heat exchanger (HX) weight.Additionally the electric power train is sized for providing the power and energy for ensuring high power ratings capability.The relevant weights are presented in Table 5.It is apparent that high effectiveness means better fuel economy.This betterment comes with the price of higher weight, longer acceleration times and reduced surge margin.As seen in Figure 21 the energy reduction achieved with the recuperated hybrid configurations is more significant compared to the simple cycle hybrid configurations for the same weight addition, even when low values of recuperator effectiveness is considered.The energy decrease per unit of added mass achieved for the recuperated configuration is five times greater compared to the simple cycle configuration.For further assessing the overall performance effect of hybridization the mission block fuel is calculated taking into consideration the increased weight of the helicopter for each configuration, as calculated.In this case the reference test case is the conventional simple cycle propulsion system for the OAG mission for TOW equal to 5000 kg.For the hybrid configurations the batteries are fully charged at the end of the mission.For the simple cycle hybrid configuration only the cases of a scaled down engine to 90% and 80% are considered since as seen in Figure 21 further increasing hybridization decrease the energy reduction per added weight.It is apparent that when the additional weight is considered the potential benefit is reduced but nevertheless the improvement is significant.As seen in Figure 22 the recuperated configurations give a significant gain in terms of block fuel despite the added weight.The betterment is in the range of 10% for a recuperator effectiveness of 40% and 18% for 80% effectiveness.
These values indicate that hybrid propulsion may be an enabling technology for recuperated cycles.The simple cycle configurations on the other hand give a marginal improvement that reduces as hybridization increases.The increase of the propulsion system weight is going to decrease the helicopter payload.If an increase in maximum TOW is selected for maintaining the maximum payload then a heavier rotor would be required.The cases of interest in this study, depicted in Figure 22 correspond to a weight increase from 4% to 10%.Another interesting aspect of hybridization is the utilization of the electric system as a buffer when transients are considered, especially for the recuperated cycles.
As seen in Figure 23 the integrated hybrid system utilizing the recuperated cycle demonstrates a rather fast throttle response since the time constants of the electrical system are rather small compared to the gas turbine [17].In this case (Figure 23) despite the throttle delay of the gas turbine, the electric power train provides the additional power to the system for the short period needed.It should be noted that gearbox inertia and potential control system lag are not considered.Additionally aspects such as circuit protection to accommodate transients should be considered as discussed in [57].

SUMMARY AND CONCLUSIONS
A parallel hybrid propulsion system utilizing simple and recuperated cycle gas

Firstly the performance and
weight of simple cycle gas turbine based configurations are assessed.The engines are scaled down versions of the baseline engine.Then the configurations utilizing the recuperated engines are assessed and their weight and performance are calculated.The different configurations are compared and the benefits accrued from utilizing hybrid propulsion based on recuperated gas turbine are assessed and discussed.

Figure 2 :
Figure 2: Simcenter Amesim engine model The gas generator consists of a five stage axial compressor (LPC), a single centrifugal stage (HPC) driven by a two stage cooled axial turbine (CT) and a two stage free power turbine (PT) delivering shaft power.The model utilizes the Simcenter Amesim readily available maps to define off-design performance for the turbomachinery

Figure 3 :
Figure3: Measured[19] versus calculated data for steady state operation Simcenter Amesim is a dynamic simulation environment; hence the model built for steady state simulation can be directly used for transient simulation.Shaft inertia is the dominant term[40].The gas generator and power turbine shaft moment of inertias are defined equal to 0.06033kg•m 2 and 0.084 kg•m 2 following the values recommended by Ballin[19].The gas dynamic effects are considered utilizing the volume components.

Figure 4 :
Figure 4: Simulated in [19] versus Simcenter Amesim calculated data for a fuel step increase As seen the model built in Simcenter Amesim and adapted to the relevant steady state data presents the expected behaviour in terms of time constants and values for transient operation as well, thus it can be used for assessing both the steady state and transient operation of the engine.

8 MCRFigure 6 :
Figure 6: Specific fuel consumption versus power for considered configurations

Figure 7 :
Figure 7: Recuperated engines sfc reduction compared to simple cycle engine ,ref[kg/s]

Figure 8 :
Figure 8: Simple cycle and recuperated configurations response to a fuel step increase

Figure 11 :
Figure 11: Simple cycle and recuperated engine response to the imposed fuel changecentrifugal compressor PROPULSION SYSTEM WEIGHT ESTIMATION
a) For the simple cycle configuration the engine design point is scaled down from 100% to 60% of the baseline engine power.The scaling is done assuming the same cycle in terms of specific power and scaling the design inlet mass flow b) For the recuperated configuration three cases assuming heat exchanger effectiveness 0.4, 0.6 and 0.8 are considered.

Figure 14 :
Figure 14: Parallel hybrid configuration for the recuperated engine

Figure 19 :
Figure 19: Overall energy and power ratio

Figure 21 :
Figure 21: Energy consumption decrease vs added weight

Figure 22 :
Figure 22: Block fuel reduction compared to conventional configuration, accounting for hybrid system weight

Figure 23 :
Figure 23: Throttle response of the integrated hybrid system and its power components: the recuperated gas turbine and the electric power train turbine configurations is assessed.Firstly, a fully adapted engine model capable to reproduce an engine steady state and transient operation is built and used as a baseline for the recuperated engine.The results indicate that the model built in Simcenter Amesim has been adapted with good accuracy to the relevant steady state data and presents the expected behaviour in terms of time constants and values when transient operation is considered.Following a recuperated version of the simple cycle engine is built and simulated for heat exchanger effectiveness ranging from 0.4 to 0.8.The expected behaviour is observed namely recuperation produces a significant betterment in terms of specific fuel consumption, especially for part load operations.Values of recuperation effectiveness in the range of 40-60%, which are considered plausible, may produce a mean sfc reduction ranging from 10% to 15% for the complete operating envelope.During transient manoeuvres the recuperated engine is lagging compared to the simple cycle due to the recuperator thermal inertia.The lagging becomes greater as the effectiveness increases.This behaviour should be assessed when recuperated engines are considered, since there are specific manoeuvres that may be affected by engine throttle lag.The addition of a heat exchanger has a detrimental effect on surge margin as well.The surge margin reduction during transient becomes more significant as the effectiveness and size of the heat exchanger increases.Having established a recuperated version of the engine a parallel hybrid configuration utilizing both the simple cycle and the recuperated configuration models is built in Simcenter Amesim.An OAG mission of a representative twin engine medium

Table 2 :
Engine ratings for the configurations

Table 3 :
Predicted and actual engine weight and length

Table 4 :
Simple cycle propulsion system weight per engine

Table 5 :
Recuperated cycle propulsion systems weight