A Multidisciplinary Approach for the Comprehensive

Ioannis Goulos Centre for Propulsion, School of Engineering, Cranfield University, Bedfordshire MK430AL, UK e-mail: [email protected]

Fakhre Ali Centre for Propulsion, School of Engineering, Cranfield University, Bedfordshire MK430AL, UK e-mail: [email protected]

Konstantinos Tzanidakis Centre for Propulsion, School of Engineering, Cranfield University, Bedfordshire MK430AL, UK e-mail: [email protected]

Vassilios Pachidis Centre for Propulsion, School of Engineering, Cranfield University, Bedfordshire MK430AL, UK e-mail: [email protected]

Roberto d’Ippolito NOESIS Solutions, Gaston Geenslaan, 11 B4, Leuven 3001, Belgium e-mail: [email protected]

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A Multidisciplinary Approach for the Comprehensive Assessment of Integrated Rotorcraft–Powerplant Systems at Mission Level This paper presents an integrated methodology for the comprehensive assessment of combined rotorcraft–powerplant systems at mission level. Analytical evaluation of existing and conceptual designs is carried out in terms of operational performance and environmental impact. The proposed approach comprises a wide-range of individual modeling theories applicable to rotorcraft flight dynamics and gas turbine engine performance. A novel, physics-based, stirred reactor model is employed for the rapid estimation of nitrogen oxides (NOx) emissions. The individual mathematical models are implemented within an elaborate numerical procedure, solving for total mission fuel consumption and associated pollutant emissions. The combined approach is applied to the comprehensive analysis of a reference twin-engine light (TEL) aircraft modeled after the Eurocopter Bo 105 helicopter, operating on representative mission scenarios. Extensive comparisons with flight test data are carried out and presented in terms of main rotor trim control angles and power requirements, along with general flight performance charts including payload-range diagrams. Predictions of total mission fuel consumption and NOx emissions are compared with estimated values provided by the Swiss Federal Office of Civil Aviation (FOCA). Good agreement is exhibited between predictions made with the physics-based stirred reactor model and experimentally measured values of NOx emission indices. The obtained results suggest that the production rates of NOx pollutant emissions are predominantly influenced by the behavior of total air inlet pressure upstream of the combustion chamber, which is affected by the employed operational procedures and the time-dependent all-up mass (AUM) of the aircraft. It is demonstrated that accurate estimation of on-board fuel supplies ahead of flight is key to improving fuel economy as well as reducing environmental impact. The proposed methodology essentially constitutes an enabling technology for the comprehensive assessment of existing and conceptual rotorcraft–powerplant systems, in terms of operational performance and environmental impact. [DOI: 10.1115/1.4028181]

Introduction

1.1 Background. Simulation of operational performance and mission analysis has always been an important topic for the rotorcraft industry. These topics are now raising even more interest as aspects related to chemical emissions and ground noise impact, gradually gain increasing importance for environmental and social impact assessments [1]. The implementation of advanced technologies for the design and production of conceptual rotorcraft, essentially entails the development of an accurate and costeffective approach applicable to the determination of performance and associated environmental impact. This is imperative when considering novel rotorcraft configurations targeting better overall fuel economy and reduced gaseous emissions. Within the operational bounds of civil aviation, current rotorcraft activities amount presently to approximately 1.5 106 flight hours per year, only with respect to the European airspace. These represent an annual consumption of the equivalent of 400,000 tons of aviation fuel [2]. The aforementioned values may seem rather insignificant when compared to the corresponding figures Contributed by the Turbomachinery Committee of ASME for publication in the JOURNAL OF ENGINEERING FOR GAS TURBINES AND POWER. Manuscript received July 4, 2014; final manuscript received July 10, 2014; published online August 26, 2014. Editor: David Wisler.

of fixed-wing aircraft traffic. However, it is emphasized that rotorcraft traffic is expected to grow significantly over the next decades, leading to associated environmental implications in terms of ground noise impact and air quality. Maintaining current design and operational technologies is expected to quadruplicate the aforementioned figures within a time-frame of approximately 20 yr. This is a direct outcome of the anticipated traffic augmentation [3]. The rotorcraft community has to respond in a pro-active manner before environmental considerations present themselves as limiting factors considering the anticipated growth of civil operations. Hence, the current objective is to come up with novel rotorcraft designs that will allow sustaining the expected growth in operations and also limit the associated environmental impact to levels considered acceptable in terms of legislation, financial, and public concerns. In light of the aforementioned background related to the design of conceptual rotorcraft, a comprehensive and simultaneously cost-effective methodology targeting the comprehensive assessment of combined rotorcraft–powerplant systems has been developed at Cranfield University. This framework has been named “HECTOR” (helicopter omnidisciplinary research) after the homonymous Trojan prince who found his tragic demise in the hands of the ancient Greek Achilles. HECTOR [4,5] aims to utilize modeling fidelity designated for rotor design applications, in the context of operational

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performance and environmental impact assessments. An advanced level of simulation fidelity is therefore incorporated in order to capture the associated performance trade-off between rotorcraft designs optimized in a multidisciplinary manner. As a result, the focus of the design process may be placed on the overall performance within designated types of operations, rather than on predefined sets of flight conditions [4]. 1.2 HECTOR Platform. HECTOR is an integrated tool that has been continuously developed over time in order to improve its modeling fidelity and expand its capabilities. Goulos et al. [5] employed HECTOR in order to investigate the potential to reduce rotorcraft mission fuel consumption, through optimization of the engine design point (DP) cycle parameters at constant technology level. The design space variables essentially comprised the combustor outlet temperature, each compressor’s pressure ratio, and the engine mass flow. A comprehensive and computationally efficient optimization strategy was deployed in order to analyze two rotorcraft classes; a TEL, and a twin-engine medium helicopter, respectively. In another study reported by Goulos et al. [4], HECTOR was employed in order to investigate the implications associated with the implicit coupling between airframe–rotor and engine performance in the context of mission analysis. Their work established that identification of the most demanding flight conditions (in terms of engine power) and their influence on total mission fuel burn cannot be addressed without catering for accurate predictions of the aircraft’s time-dependent AUM during flight. This essentially signified the necessity for accurate estimation of the interrelationship between engine and airframe–rotor performance throughout the course of the designated operation, due to their implicit coupling through the aircraft AUM. The theoretical and computational development of HECTOR has been documented extensively in Ref. [6]. 1.3 Modeling of Helicopter Gaseous Emissions. Gaseous emissions of gas turbines engines are essentially combustion products of aviation fuel with air. Aviation fuel consists of approximately 99.7% of hydrocarbons, represented as CxHy, and up to 0.5% of sulfur. Combustion products comprise primarily, among others, chemical equilibrium emissions such as CO2 and H2O, as well as trace levels of sulfur oxides (SOx). Incomplete combustion leads to carbon monoxide (CO) and unburned hydrocarbons (UHC). A portion of the UHC emissions include trace amounts of toxic species referred to as hazardous air pollutants. Additional carbon is emitted in the form of soot and smoke. Due to the high temperatures of combustion and the presence of nitrogen (N2) in the intake air, nitric and NOx are also emitted. CO2 and H2O gaseous emissions are products of chemical equilibrium and can be reduced by developing more efficient engines, e.g., regenerated engines or high overall pressure ratio (OPR) engines. In the short term (25 yr or less), reduction in CO2 and H2O emissions will most likely be achieved by improving fuel efficiency. This is due to the fact that, employment of alternate fuels requires major changes in aircraft and engine design, along with the development of infrastructure for proper delivery and storage. Pollutants such as UHC, CO, and smoke are products of incomplete combustion and can be reduced by improvements in combustor design. Mitigation of NOx emissions is targeted by achieving thermal uniformity within the combustor chamber, mainly by limiting peak temperatures and reducing residence times in high-temperature regions of the flame. Gaseous emissions of helicopter gas turbine engines are not easily assessed (specifically NOx) due to lack of available emissions data for turboshaft engines in the public domain. It is also emphasized that there is no generally recognized approach on how to estimate gaseous emissions of helicopter engines (HELEN). For this reason, in 2008, the Swiss FOCA launched a project named HELEN [7]. Through the HELEN project, an 012603-2 / Vol. 137, JANUARY 2015

empirical approach was employed and measurements of turboshaft engine emissions were made during tests carried out after overhaul. As a result, empirical functions for helicopter engine emission factors were derived based on systematic experimental test measurements conducted on a wide-range of helicopter turboshaft engines. 1.4 Scope of Present Work. In light of the research already presented in the existing literature, it can be realized that HECTOR has been utilized extensively for the comprehensive assessment of integrated rotorcraft–powerplant systems in terms of flight dynamics, airframe–rotor design, and engine performance. The associated environmental impact along with the corresponding implications in the context of mission analysis, have not yet been evaluated or addressed in the literature. This work aims toward closing this gap with the development and implementation of an innovative numerical formulation for the rapid estimation of gaseous emissions, predominantly considering the formation of NOx. A physics-based stirred reactor model is employed, in which the turbulent flow within the various combustor zones is sufficiently idealized and the time-dependent chemistry governing the formation of the pollutant species of interest is accounted for. The proposed method is based on first principles and is applicable to traditional combustor designs incorporated by helicopter turboshaft engines. The combined approach is deployed for the comprehensive analysis of a reference TEL rotorcraft, modeled after the Eurocopter Bo 105 helicopter. Assessments are performed for two representative operations defined in collaboration with the European Helicopter Operators Committee (EHOC); a passenger air taxi mission (PATM), and a law enforcement mission (LEM). Extensive comparisons with flight test data on the main rotor trim controls and power requirements is presented along with calculated payload–range diagrams. Predictions of total mission fuel consumption and gaseous NOx emissions are compared against estimations carried out using the closed form expressions established by FOCA [7]. The time-variations of certain engine performance parameters of interest, along with the behavior of the produced NOx, are investigated for both rotorcraft operations. Good agreement is found between predictions made with the physics-based stirred reactor model and experimentally measured values of NOx emission indices. It is shown that the production rate of NOx emissions during a typical TEL rotorcraft mission is affected predominantly by the behavior of total air inlet pressure upstream of the combustion chamber. It is argued that the production rate of NOx is influenced significantly by the employed flight conditions and the aircraft AUM. Finally, it is demonstrated that accurate estimation of on-board fuel supplies ahead of flight, not only has the potential to improve fuel economy, but is also key to reducing the associated environmental impact. The proposed approach constitutes an enabler for the comprehensive evaluation of conceptual rotorcraft at mission level, in terms of operational performance and environmental impact.

2

Simulation Methodology

2.1 Mission Performance Simulation of Integrated Rotorcraft–Powerplant Systems (HECTOR). HECTOR comprises a finite series of consecutive analyses, each applicable to a different aspect of rotorcraft flight dynamics and engine performance. HECTOR constitutes an integrated approach that comprises the Lagrangian rotor blade modal analysis method developed in Refs. [6] and [8], a flight path profile analysis based on the World Geodetic System dated in 1984 (WGS 84) [9], the nonlinear trim procedure solving for the aeroelastic behavior of the main rotor blades developed in Refs. [6], [10], and [11], the engine offdesign (OD) performance analysis reported in Ref. [12], and the gas turbine emissions prediction model presented in Ref. [13]. The individual modeling methodologies are elaborately integrated Transactions of the ASME


within an overall numerical procedure, solving for the unknown initial aircraft AUM. A detailed description of HECTOR has been provided by Goulos et al. in Refs. [4–6]. 2.2 Gas Turbine Performance Simulation (TURBOMATCH). The engine performance model (TURBOMATCH) employed for the simulations carried out in this paper has been developed over a number of decades at Cranfield University [12]. Turbomatch is based on zero-dimensional analysis of the aerothermodynamic processes occurring throughout the engine gas path, employing discrete component maps. The incorporated methodology essentially solves for the mass and energy balance between the various engine components. TURBOMATCH has been previously deployed in several studies available in the literature for the prediction of DP and OD performance of gas turbine engines [14,15]. In order to comply with the scope of work presented in this paper, the engine is assumed to be operating exclusively at steady-state OD conditions during the overall mission course. 2.3 Prediction of Gaseous Emissions (HEPHAESTUS). In order to predict the gaseous emissions arising from fossil fuel combustion, the deployment of a robust prediction methodology is necessary. Thus, a generic emission indices prediction software has been developed at Cranfield University under the name of HEPHAESTUS [13,16]. HEPHAESTUS constitutes a general prediction methodology based on the stirred reactor concept along with a set of simplified chemical reactions. HEPHAESTUS is capable of accounting for the design specifications of any modeled combustion system. Thus, the user can specify a combustor geometry in terms of primary, intermediate, and dilution zone volumes as well as the mass flow distribution of a given combustor design. HEPHAESTUS has been previously adopted in several aircraft trajectory optimization studies, e.g., Ref. [17]. 2.4 Integration of HEPHAESTUS Within HECTOR. Figure 1 displays an illustrative flow-chart that describes the numerical approach employed in HECTOR, for the simulation of complete rotorcraft operations. The user-defined mission is essentially subdivided within designated mission task elements. Examples of mission task elements include: hover, forward flight, climb, descent, take-off, and land. Each mission task element is broken down into discrete segments with a predefined time-step Dt according to user-specification. The helicopter is assumed to be operating in trim and the engine in steady-state OD conditions during each segment. This assumption is deemed valid considering the fact that, that the dominant focus of the described approach is the evaluation of total mission fuel consumption along with the associated gaseous emissions. It is demonstrated in Fig. 1 that the overall simulation process is initiated (N ¼ 1) with an initial guess of the on-board fuel supplies. Depending on the user-defined operation and the corresponding segment time-point t, the flight path model determines the respective flight conditions. The nonlinear flight dynamics model (HECTOR) subsequently calculates the rotorcraft trim state for the imposed set of flight conditions. This process essentially determines the engine shaft power requirement along with the corresponding inlet conditions at the intake. The engine performance model (TURBOMATCH) then establishes the engine operating point, thus determining the engine fuel flow and the combustor inlet conditions. This data are then utilized by the combustor emissions model (HEPHAESTUS), which determines the emission indices for the particular species of interest. The timedependent fuel consumption at time t along with the respective emissions inventory is then updated by applying a numerical time integration scheme on the time-variations of engine fuel flow and emission indices for 0 s t. The calculated value of fuel burn is then subtracted from the initial AUM in order to account for the gradual weight reduction during the course of the mission. The Journal of Engineering for Gas Turbines and Power

Fig. 1 Illustration of the numerical formulation employed in HECTOR for the simulation of complete rotorcraft operations [4]

overall process is re-iterated in a fixed-point manner until convergence is obtained for the total mission fuel burn. 2.5 Compilation of Integrated Rotorcraft–Powerplant System Configuration. The aircraft configuration assumed for the purpose of this study has been modeled after the Eurocopter Bo 105 helicopter. The Bo 105 is a TEL utility multipurpose helicopter, equipped with two Rolls-Royce Allison 250-C20B turboshaft engines rated at 313 kW maximum contingency shaft power. The Allison 250-C20B engine is equipped with a single-spool gas generator including a six-stage axial followed by a centrifugal compressor. Table 1 presents the rotorcraft model characteristics along with the corresponding powerplant configuration. The maximum contingency power setting has been selected as the DP for JANUARY 2015, Vol. 137 / 012603-3


Table 1 Baseline design parameters for the reference Bo 105 rotorcraft–powerplant system Design parameter

Value

Unit

Bo 105 type rotorcraft (airframe and rotor) OW MTOW Number of blades Blade radius Blade chord Blade twist Rotorspeed

2200 2500 4 4.91 0.27 8 2544

kg kg — m m deg deg/s

Allison 250-C20B type powerplant TET W LPC PR HPC PR DP shaft power DP SFC

1470 1.56 2.73 2.6 417 109.98

K kg/s — — hp lg/J

the respective TURBOMATCH model. The model has been matched at DP conditions in terms of specific fuel consumption (SFC), with public domain data extracted from Ref. [18]. The configuration of the Bo 105 as well as its performance characteristic have been extensively documented and analyzed in Ref. [19]. A detailed description of the Allison 250-C20B engine family can be found in Ref. [18]. 2.6 Reference Combustor Modeling. As an input requirement to the emission model (HEPHAESTUS) the combustor geometry of a reference Rolls-Royce Allison 250-C20B engine was investigated in detail. A “reverse engineering” approach was adopted by means of publicly available data. The complete, fullscale, three-dimensional combustion chamber was modeled using CATIA Part Modeling to derive the best possible approximations for the inlet area, outlet area, volume, and length of the various “zones” (e.g., flame front primary, intermediate, and dilution zones) within the combustor. The Rolls-Royce Allison 250-C20B engine is equipped with a straight through single can combustor module [18]. The coordinates of the combustor module were obtained from two-dimensional engine cutaway drawings available in the public domain. The obtained coordinates were subsequently exported to CATIA Part Modeling in order to obtain a digital three-dimensional representation for the reference combustor module, as shown in Fig. 2. The obtained digital combustor design was then used to approximate the dimensions of the different zones (lengths and volumes) within the corresponding model in HEPHAESTUS. The combustion chamber is represented by combining four distinctive zones, all simulated using a series of perfectly stirred reactor (PSR) models, as shown in Fig. 3. The theory behind PSR

Fig. 2 Geometric model of the annular combustion chamber of the reference Rolls-Royce Allison 250-C20B engine

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Fig. 3 Combustor zones, fuel/air fractions, and reactor models assumed for the modeling of gaseous emissions

focuses predominantly on the influence of the different levels of turbulent mixing on the chemical reactions within prescribed volumes, rather than on the details of spatially varying velocities and turbulence fields within a combustor. It is also assumed that the flame front is well mixed and can also be approximated with a PSR model. Each reactor assumes a chemical equilibrium state for calculating the reactor temperature, pressure, and density. The model includes 18 species for the equilibrium calculations using the NASA CEA program [20]. The NOx mass fraction variation is calculated using the extended Zeldovich mechanism (includes N2O mechanism) and Prompt NOx methodology reported in Refs. [17] and [21–23]. The analysis assumes that, 90% of the fuel is burned at the flame front zone at a fixed equivalent ratio of 0.9. The remaining fuel fraction is assumed to be burned within the region of the combustor primary zone. The distribution of air mass flow fractions within each of the combustor zones is in accordance with sensible engineering estimations and it is assumed to remain constant at OD operating conditions. The volume of each reactor is estimated based on the combustor geometry information obtained from the developed CATIA model.

3

Results and Discussion

3.1 Reference Combustor NOx Prediction. Accurate determination of NOx gaseous emissions from helicopter gas turbine engine combustors can quickly turn into a time consuming and laborious exercise. This is due to the complicated nature of the physical processes occurring in the combustion chamber. The methodology adopted in this paper attempts to capture some of the dominant physical phenomena taking place in an actual combustor using first principles. It is emphasized that the fuel fractions, air fractions, and different combustion zones split are estimated based on engineering judgment along with publicly available information. This is in accordance with previously reported studies [13,24]. Figure 4 demonstrates the predicted variation of the NOx emission index with equivalence ratio. The concentration of NOx reaches the highest level near the stoichiometric mixture, on the lean side. The total NOx formed in the combustor is the summation of the concentration arising from each reactor. As such, the NOx formation rate is largely dependent on the corresponding conditions within each reactor. Increasing equivalence ratio essentially raises gas temperature, which inevitably increases the formation of thermal NOx and therefore increases the total production of NOx. This effect occurs only until the stoichiometric equivalence ratio. Beyond this point, the total amount of fuel in the mixture is essentially deprived of available oxygen and thus, complete combustion is unattainable. Figure 5 compares HEPHAESTUS predictions for the variation of NOx emission index against engine shaft power, with experimental data as well as with the empirical approximate function derived and reported by FOCA in Ref. [7]. Experimental data are included considering two families of engines for which FOCA conducted NOx measurements. These effectively represent the Transactions of the ASME


Fig. 4 Flame front equilibrium temperature and NOx emission index as functions of equivalence ratio for the reference RollsRoyce 250-C20B combustor model

lower and upper bounds of the overall deviation observed in the measured data. The behavior of the approximate FOCA expression is also included in Fig. 5 for completeness. HEPHAESTUS has been matched at DP conditions considering both the lower and upper bounds of the measured data, as well as with the approximate function proposed by FOCA. Reasonable agreement can be observed throughout the power range for which results are presented, considering all three cases. Specifically, the trend of the NOx emission index with engine power has been captured adequately by the proposed approach. It is emphasized that the combustor zone sizing process was essentially carried out with the objective to align the predicted values of NOx against the data corresponding to each case, at DP conditions (417 hp) only. Once the predicted value of NOx emission index was well matched against the corresponding data at DP, the combustor details were kept fixed throughout the OD power range. The deviation observed in Fig. 5 between HEPHAESTUS predictions and the FOCA data considering lower engine power settings are attributed predominantly to OD engine performance modeling deficiencies associated with the use of generic and scalable component maps in TURBOMATCH [12]. 3.2 Helicopter Flight Dynamics Trim Analysis. Figure 6 presents rotor trim controls and power requirement predictions obtained from nonlinear trim simulations carried out with the nonlinear flight dynamics model of HECTOR [4,10,11]. Results are presented for straight and level flight as functions of advance ratio ðl ¼ V=XRÞ, from hover (l ¼ 0) to high-speed flight (l 0.36).

Fig. 5 NOx emission index prediction for the reference RollsRoyce Allison 250-C20B engine—comparison against closed form approximations and experimental measurements derived by FOCA [7], including representative lower and upper bounds

Journal of Engineering for Gas Turbines and Power

Fig. 6 Trim performance predictions for the Bo 105 rotorcraft—comparison with flight test data from Ref. [25]: (a) rotor power required, (b) collective pitch angle h0, (c) longitudinal cyclic pitch angle h1s, and (d) lateral cyclic pitch angle h1c

Comparisons with flight test data extracted from Ref. [25] are also included for validation purposes. Figures 6(a) and 6(b) present trim values of main rotor power requirement and collective pitch angle (h0), respectively. Very good correlation between the integrated model predictions and flight test data can be observed regarding both trim outputs, especially considering low values of advance ratio. However, the agreement is not as good with respect to higher values of advance ratio. This is attributed to possible deficiencies in the employed aerofoil data, along with the underprediction of aerodynamic fuselage drag in the look-up tables used in HECTOR [6]. Figures 6(c) and 6(d) present trim values of longitudinal (h1s) and lateral (h1c) cyclic pitch angles, respectively. Good agreement is observed between predictions made with the present model and flight tests, considering both cyclic input controls. Rotor wake impingement on the horizontal stabilizer leads to underprediction of h1s for l 0.02. h1s is also underpredicted for l 0.15 due to the aforementioned underprediction of airframe drag and the associated aerodynamic pitching moment. A noticeable deviation between predicted and measured lateral cyclic (h1c) for l 0.05, reveals the requirement for more sophisticated wake modeling considering the specific flight region. 3.3 Integrated Rotorcraft–Powerplant Engine Performance. Figures 7(a)–7(d) present the respective engine performance parameters as functions of cruise speed. Results are JANUARY 2015, Vol. 137 / 012603-5


It is emphasized at this point that ram temperature rise may increase the compressor work for a given pressure ratio. However, with regards to the designated operating conditions and engine configuration, the former effect is dominant, leading to a reduction in the power absorbed from the gas generator turbine for compression. The corresponding increase in air mass flow due to ram compression is of the order of 0.98%. The aforementioned effects essentially lead to a reduction in engine fuel flow of the order of 2.6% for a cruise speed of 225 km/h in comparison to the low cruise speed case 30 km/h, while producing almost the same value of shaft power Pengine. Figures 7(c) and 7(d) present the respective correlations acquired for the reference Bo 105 helicopter in terms of engine power, fuel flow, and NOx production rate. It can be observed that the production rate of NOx also varies in a trend that is quite similar with the engine shaft power. However, the NOx rate for a given level of engine power increases with cruise speed. This is also attributed to the effect of Mach number, which increases combustor inlet pressure, and thus the emission index of NOx.

Fig. 7 Engine performance parameters during trimmed flight for the reference Bo 105 rotorcraft: (a) engine fuel flow Wf and shaft power Pengine, (b) SFC and TET, (c) NOx production rate and engine shaft power Pengine, and (d) engine fuel flow Wf and NOx production rate

presented for atmospheric conditions corresponding to zero international standard atmosphere (ISA) deviation at an absolute altitude of 100 m. The intake Mach number Mintake is essentially defined by the designated cruise speed. It can be observed that the results obtained for the presented engine performance parameters essentially follow the classic “bucket-shape” curve, which is dictated by the power requirement of the main rotor, as shown in Fig. 6(a). It is noted that a small ram compression effect on engine performance is present in the results presented in Figs. 7(a) and 7(b). Figure 7(a) demonstrates that the required engine shaft power Pengine is roughly 180 kW for a cruise speed of approximately 30 km/h, as well as for 225 km/h. The latter cruise speed value corresponds to Mintake 0.2, which essentially leads to an increase of total inlet total pressure of the order of 2.7% due to ram compression. The predicted change in maximum engine pressure is around 0.5% upwards. However, the OPR produced solely by the compressors is reduced by approximately 2.1%, leading to a corresponding decrease in compressor work. 012603-6 / Vol. 137, JANUARY 2015

3.4 Derivation of Performance Charts and Payload-Range Diagrams. In order to construct the required payload-range diagrams with regards to the rotorcraft employed for the investigations presented in this paper, the relationship between specific range and operational weight (OW) needs to be established. The maximum specific range is obtained at a flight speed known as velocity for best range (Vbr). This can be estimated by drawing a tangent line to the fuel flow curve (Fig. 7(a)), starting from the axes origin [26]. It is noted that Vbr is significantly influenced by the rotorcraft’s OW and that it increases with increasing rotorcraft weight as well as altitude [26]. Thus, HECTOR was employed in order to obtain the corresponding fuel flow curves for different values of OW. A number of nonlinear trim performance simulations was carried out starting with zero payload (OW ¼ operational empty weight (OEW)), and subsequently increasing the rotorcraft’s OW gradually up to the maximum take-off weight (MTOW). The operational empty weight simply dictates zero useful payload, while maximum payload essentially corresponds to AUM ¼ MTOW. Simulations were carried out for sea-level standard ISA conditions. Figure 8 presents the rotorcraft’s specific range as a function of flight speed for OW ¼ MTOW along with the corresponding fuel flow curve. It is noted that another important value of flight speed is that corresponding to minimum engine fuel flow. This is essentially known as velocity for best endurance (Vbe). At this speed, the power required for forward flight is minimum. Thus, the excess engine power available is maximum, and therefore, the maximum rate of climb can be accomplished [26]. The payload-range diagram shown in Fig. 9 is constructed on the basis that the helicopter mission consists of a 5-min idle (warm up) period followed by a take-off maneuver and climbing flight up to an altitude of 200 m. After climbing, the rotorcraft is assumed to cruise at the Vbr corresponding to the average aircraft AUM, and subsequently descents and lands with 0% fuel reserves. After that, a 1-min long idle period is assumed at 20% of total engine power.

Fig. 8 Reference Bo 105 rotorcraft performance estimation: specific air range at MTOW—cruise at 200 m altitude

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Fig. 9 Reference Bo 105 rotorcraft performance estimation: payload-range diagram

3.5 Compilation of Representative Mission Scenarios. Two generic three-dimensional reference missions, representative of modern TEL rotorcraft operations, were designed; a PATM and a LEM. The incorporated operational procedures in terms of geographical location, airspeed, altitude, climb/descent rates, and idle times have been defined in collaboration with European Helicopter Operators’ Committee (EHOC). The geographical representation in terms of global coordinates, along with the employed operational procedures, is illustrated in Fig. 10 for the PATM and Fig. 11 for the LEM, respectively. The PATM assumes that the rotorcraft takes-off from Hanover Airport (Germany) and travels in order to pick up passengers from a designated location. It subsequently transfers them to Garbsen Hotel and then transits back to Hanover Airport. The LEM schedule assumes that the helicopter takes-off from Myttinge heliport (Stockholm) and subsequently conducts high altitude surveillance over two different areas. Once the surveillance protocol is complete, the rotorcraft transits back to Myttinge Heliport, where it originated from. The implemented operational procedures with respect to both missions were defined within the typical operational envelope encountered by modern TEL rotorcraft. Climb and descent rates were held fixed at 5 m/s considering both missions. A time-step Dt of 5 s was used for each individual mission segment. All coordinated turns were assumed to be executed with a turn rate of 5 deg/s. During idle operation, the overall helicopter power

Fig. 10 Reference PATM: (a) geographical definition and (b) time-variations of deployed operational airspeed and AGL altitude


Fig. 11 Reference LEM: (a) geographical definition and (b) time-variations of deployed operational airspeed and AGL altitude

requirements are assumed to be equal to 20% of DP (maximum contingency) engine shaft power. Figures 10(b) and 11(b) present the time-variations of above ground level (AGL) altitude and flight speed for the PATM and LEM, respectively. With regards to the PATM, it can be noticed from Fig. 10(a) that the mission comprises three legs in total. The cruise speed within the first two legs is 60 m/s at an altitude of 450 m, while for the final leg, the cruise speed is 50 m/s at an altitude of 350 m. With respect to the LEM, Fig. 11(b) shows that the helicopter spends the majority of the mission time engaged in loitering (surveillance), where the deployed flight speed and AGL altitude are 40 m/s and 300 m, respectively.

3.6 Operational Performance and Environmental Impact Assessment. HECTOR has been applied for the evaluation of operational performance and environmental impact of the Eurocopter Bo 105 rotorcraft–powerplant system (Table 1) with respect to the PATM and LEM operations. The total mission parameters in terms of time, range, fuel consumption, and gaseous emissions are outlined in Table 2 considering both helicopter operations. Separate simulations have been carried out using the HEPHAESTUS models matched to all three cases of NOx inventories extracted from the FOCA report [7], as shown in Fig. 5 (lower and upper experimental bounds, and FOCA closed form approximation). It can be noted that the prediction of total NOx for the cases that HEPHAESTUS is matched to the lower and upper bounds of the FOCA data, deviates from the approximative case (FOCA approx.) by roughly 57% and 28%, respectively, considering both rotorcraft operations. It is important to notice that the aforementioned deviations are identical for both missions. This essentially indicates the validity, consistency, and robustness of the nonlinear approach employed in HECTOR. The fact that the deviation relative to the FOCA approximation is larger considering the lower data bound (57%) can be easily explained by inspecting the results presented in Fig. 5. Specifically, it is noted that the FOCA expression is consistently closer to the upper bound JANUARY 2015, Vol. 137 / 012603-7


Table 2 Reference PATM and LEM parameters for the reference Bo 105 rotorcraft–powerplant system Mission parameters

PATM

LEM

Unit

Time Range Fuel burn CO2 H2O NOx FOCA approx.

1725 36.22 60.04 191.92 74.6 0.287

3945 162.76 188.03 600.10 233.5 0.981

s km kg kg kg kg

NOx deviations (%) NOx lower bound NOx upper bound

57.14% 27.87%

56.88% 27.93%

Rel. to FOCA approx. Rel. to FOCA approx.

throughout the majority of the power range considered, compared to the respective lower bound of the measured data. Table 3 compares nonlinear predictions made with HECTOR for total mission fuel consumption and NOx emissions, with estimates derived by exploiting certain simplified expressions, also suggested by FOCA in Ref. [7]. According to FOCA, the amount of total mission fuel burn Fb can be approximated by simply multiplying the mean mission fuel flow Wf,mean with the amount of time spent in operation. This process is described by Eq. (1a). The emissions inventory for NOx can also be established in a similar fashion by multiplying the estimated value of fuel burn with the mean Þ, see Eq. respective NOx emission index for mean power ðEIPNO x (1b). The corresponding predictions for CO2 and H2O are not compared with the correlations provided by FOCA. This is because the respective emission indices are almost constant with engine shaft power due to the fact that CO2 and H2O are essentially products of chemical equilibrium Fb1h ¼ 3600 Wf;mean Nengines NOx1h ¼ 3600 Wf;mean

mean EIPNO x

(1a)

Nengines

(1b)

Equations (1a) and (1b) were utilized in order to establish estimates for fuel burn and NOx gaseous emissions considering both rotorcraft operations. The estimates established are presented in Table 3. It can be observed that, for fuel burn, a relative error of less than 1% is achieved as regards both rotorcraft operations. However, considering the emissions inventory of NOx, the relative error ranges from roughly 9.1% to almost 18.7% for the PATM, and from 3.6% to around 6% for the LEM. The actual error values observed essentially depend on the FOCA dataset with which HEPHAESTUS is matched, as shown in Fig. 5. The higher relative error observed for the PATM when HEPHAESTUS is matched to the FOCA approximation, is

attributed to two specific elements. The respective Pmean 167 hp and operating power at idle Pidle 41 hp for PATM, both fall under the low power end as shown in Fig. 5. Thus, the agreement between the FOCA data and HEPHAESTUS has a discrepancy of around 8% at Pmean and around 25% for Pidle. This essentially leads to a higher relative error in the estimates for total produced NOx, due to the fact that the amount of time spent idling during the PATM is approximately 41% of total mission time. With regards to the LEM and the approximate FOCA equation, the relative error for the overall produced NOx is 6%, when compared to estimates derived using the simplistic expressions suggested by FOCA. This is due to the fact that Pmean for the LEM is approximately 254 hp. Figure 5 shows that the agreement between the FOCA expression and HEPHAESTUS predictions for the corresponding power setting is excellent since no relative deviation is observed. The observed value of 6% in relative error is introduced due to mission time spent at low power mission task elements, such as idling and descending flight. The lower values of relative error observed in Table 3 when HEPHAESTUS is tuned to the lower and upper bounds of the measured FOCA data are attributed to the relatively better alignment of the emissions model with the respective data points. This behavior is evident from the results presented in Fig. 5. Figures 12–15 present charts for various engine performance parameters of interest such as: shaft power, fuel flow, turbine entry temperature (TET), and SFC. The rates at which gaseous emissions are formed, such as: CO2 and NOx are also shown as function of mission time. In order to avoid repetition, NOx production rates are only shown for the case that HEPHAESTUS is matched to the approximate FOCA expression. Although the absolute values of the NOx formation rates differ when HEPHAESTUS is tuned to different datasets, the corresponding trends are essentially identical. Results are presented for both investigated operations. The idle segments can be distinguished as conditions associated with low values of engine shaft power, fuel flow, CO2, NOx, and TET but relatively high SFC. Climbing forward flight along with hover are identified as the most demanding trim settings in terms of engine shaft power. As already established by Goulos et al. in Ref. [4], the relationship between hover and climbing forward flight in terms of rotor power requirement depends explicitly on the timedependent AUM of the aircraft. Reference [4] showed that a large AUM may impose a more significant penalty on rotor power

Table 3 Comparison of predicted gaseous emissions inventories using nonlinear HECTOR simulations and simplified approximations based on mean engine performance parameters Mission parameters Pmean Wf,mean mean EIPNO FOCA approx. x mean EIPNO lower bound x mean EIPNO upper bound x Relative errors (RE) (%) REfb RENOx FOCA approx. RENOx lower bound RENOx upper bound

PATM

LEM

Unit

167 0.0175 3.86 1.85 5.32

254 0.0238 4.91 2.17 6.29

hp kg/s g/kg of fuel g/kg of fuel g/kg of fuel

0.55 18.78 9.19 12.48

0.13 6.01 3.66 5.88

% % % %

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Fig. 12 Engine performance parameters for the PATM: (a) shaft power Pengine—fuel flow Wf and (b) shaft power Pengine—NOx production rate



Fig. 13 Engine performance parameters for the PATM: (a) shaft power Pengine—CO2 production rate and (b) TET—SFC

Fig. 15 Engine performance parameters for the LEM: (a) shaft power Pengine—CO2 production rate and (b) TET—SFC

requirement in hover rather than climbing or forward flight. This is due to the associated reduction in mean induced flow during forward flight conditions. This analysis can now extrapolated in order to investigate the influence of the aforementioned behavior on the production rates of gaseous emissions, and specifically on the production of NOx. This behavior is shown explicitly in Figs. 12(a) and 12(b) with respect to the PATM. The power requirement, fuel flow, and the production rate for NOx considering the first hover task element ð200s t 300sÞ, are all lower compared to climbing forward flight taking place immediately afterwards ð300s t 400sÞ. As regards the second hover task element however ð950s t 1050sÞ, it can be observed that the corresponding performance parameters and production rates for gaseous emissions are almost identical for hovering and climbing forward flight ð1050s t 1200sÞ immediately afterwards. This is due to the fact that for the second hover task element, the helicopter has picked-up a number of passengers. This has essentially increased

the aircraft AUM considerably, leading to higher requirement for rotor thrust and therefore power. In accordance with the findings reported by Goulos et al. in Ref. [4], an interesting behavior is once again observed with respect to climbing/descending forward flight, considering both designated operations. Figures 12(a) and 14(a) show that during climb there is a gradual increase in engine shaft power. This is due to the gradual decrease in air density, leading to increased induced flow through the main rotor disk for a given thrust setting. This results in a corresponding rise in induced losses. However, this is not the case with respect to fuel flow. Fuel flow is actually shown to decrease with altitude, despite the rise in engine shaft power (a separate plot of climb segment is presented in Fig. 16(a) for shaft power against fuel flow). The drop in fuel flow is attributed to the gradually decreasing ambient temperature, which gives rise to an increased referred rotational speed. The results presented in Fig. 16(c) suggest an increase in OPR and TET that combine to increase the engine’s thermal efficiency. This effectively outweighs the increase in power requirement. With regard to the production rate of NOx, it is emphasized that the respective emission index is very sensitive to the combustor air inlet conditions, them being total pressure, total temperature, and mass flow. The production rate of NOx is also quite sensitive with respect to other combustor related parameters including engine fuel flow, equivalence ratio, equilibrium peak temperatures, and residence times. For the purpose of this work, the equivalence ratio was held fixed and close to the stoichiometric equivalence ratio of 0.9, as shown in Fig. 4. Thus, the following analysis essentially focuses only on the combustor inlet conditions in terms of total temperature and pressure. Any change in inlet air mass flow is adjusted accordingly in order to achieve a stoichiometric equivalence ratio equal to 0.9. The obtained results suggest that the combustor inlet temperature does not fluctuate and remains almost constant, as presented in Fig. 17(b). Thus, the production rate of NOx is seen to be mostly affected by the variation in combustor inlet pressure that is shown in Fig. 17(a). It is noted that a reduction of roughly 2.37% is observed in the combustor inlet pressure which is caused by the associated drop in ambient pressure with increasing altitude. This behavior essentially causes a decrease of the order of 5% in the production rate of NOx considering the specific conditions corresponding to the given climb segment. Reference [4] elaborated on the magnitude of the rotorcraft’s gradual weight reduction effect due to the associated fuel

Fig. 14 Engine performance parameters for the LEM: (a) shaft power Pengine—fuel flow Wf and (b) shaft power Pengine—NOx production rate


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was therefore emphasized. In the context of this work, it is understood that overestimation of the on-board fuel supplies may not only lead to a penalty in fuel economy, but also in the excessive production of associated pollutant emissions. However, underestimation of the on-board fuel supplies may essentially render a designated mission unfeasible as demonstrated in Ref. [4]. It can thus be concluded that accurate prediction of the overall mission fuel consumption ahead of flight is not only a key parameter in addressing the aspect of fuel economy, but also toward limiting the production of gaseous emissions.

4

Fig. 16 Engine performance parameters during climbing flight for the PATM: (a) shaft power Pengine—fuel flow Wf, (b) fuel flow Wf—CO2 production rate, and (c) OPR—TET

consumption, on the engine shaft power requirement and the associated performance parameters. The necessity for accurate estimation of the inter-relationship between engine and airframe–rotor performance in the context of mission analysis, due to their implicit coupling through the aircraft’s time-dependent AUM,

Conclusions

This paper has described an integrated approach for the comprehensive assessment of combined rotorcraft–powerplant systems, in terms of operational performance and environmental impact at mission level. A physics-based stirred reactor model targeting the rapid estimation of gaseous emissions has been described and validated. The developed methodology has been based on first principles and is suitable for modeling traditional combustor designs incorporated by helicopter turboshaft engines. The proposed approach has been implemented in a multidisciplinary simulation framework for the analytical evaluation of existing and conceptual rotorcraft. The combined methodology has been deployed for the comprehensive analysis of a generic TEL helicopter, operating within two representative missions. Extensive comparisons have been carried out with flight test data on main rotor trim controls and power requirements. The corresponding payload-range diagrams have been calculated. Nonlinear predictions of total mission fuel consumption and gaseous NOx emissions have been compared against experimental measurements, as well as with approximate closed form expressions established by FOCA. The time-variations of certain engine performance parameters, along with the behavior of gaseous emissions, have been investigated. Reasonable agreement has been found between predictions made with the proposed physics-based stirred reactor model and experimentally measured values of NOx emission indices. Good correlation has also been found in terms of total mission fuel burn and NOx production, between nonlinear simulations and simplified expressions provided by FOCA. The obtained results have shown that for the typical operational envelope encountered by modern TEL rotorcraft, the favorable altitude effect on engine thermal efficiency dominates over the respective penalizing influence on airframe–rotor performance. It has been demonstrated that under these operating conditions, the production rate of NOx emissions is affected predominantly by the behavior of total air inlet pressure upstream of the combustion chamber. It has been argued that the production rate of NOx pollutants is significantly influenced by the employed flight conditions as well as by the time-dependent aircraft AUM. Finally, it has been demonstrated that accurate estimation of on-board fuel supplies ahead of flight, not only has the potential to improve fuel economy, but is also key to reducing the associated environmental impact. The proposed methodology constitutes an enabling technology for the evaluation of existing and conceptual rotorcraft, in terms of operational performance and environmental impact at mission level.

Acknowledgment The authors would like to acknowledge Professor Pericles Pilidis, Dr. Vishal Sethi, Dr. Hugo Pervier, and Mr. Atma Prakash from the department of Power and Propulsion of Cranfield University, for their insightful advice and continuing support.

Nomenclature Fig. 17 Engine performance parameters during climbing flight for the PATM (climb rate 5 5 m/s): (a) combustor inlet pressure P3—NOx production rate and (b) combustor inlet temperature T3—NOx production rate

012603-10 / Vol. 137, JANUARY 2015

mean EIPNO ¼ NOx emission index for mean engine shaft power x (g/kg of fuel burn) Fb ¼ fuel burn (kg) Fn ¼ fuel burn calculated for the nth mission iteration (kg)



Mintake ¼ n¼ Nengines ¼ Pengine ¼ Pidle ¼ Pmean ¼ P3 ¼ R¼ REfb ¼ T3 ¼ Vbe ¼ Vbr ¼ Wf ¼ Wf,mean ¼ W3 ¼

engine intake Mach number fuel burn iteration index number of engines engine shaft power (hp/kW) engine idle shaft power (hp/kW) engine mean shaft power (hp/kW) combustor air inlet pressure (atm) helicopter rotor blade radius (m) relative error for mission fuel burn RENOx relative error for emissions inventory of NOx combustor air inlet temperature (K) airspeed for maximum endurance (m/s) Airspeed for maximum range (m/s) engine fuel flow (kg/s) engine mean fuel flow (kg/s) combustor air inlet mass flow (kg/s)

Greek Symbols Dt ¼ e¼ h0, h1c, h1s ¼ l¼ X¼

mission analysis time-step (s) mission fuel consumption tolerance main rotor trim control angles (deg) advance ratio, V=XR main rotor nominal rotorspeed (deg/s)

Acronyms ACARE ¼ Advisory Council for Aeronautics Research in Europe AGL ¼ above ground level AUM ¼ all-up mass (kg) CATIA ¼ computer aided three-dimensional interactive application DP ¼ design point EHOC ¼ European Helicopter Operators’ Committee FOCA ¼ Federal Office of Civil Aviation HECTOR ¼ helicopter omnidisciplinary research-platform HELEN ¼ helicopter engines HPC ¼ high-pressure compressor ISA ¼ international standard atmosphere LEM ¼ law enforcement mission LPC ¼ low-pressure compressor MTOW ¼ maximum take-off weight (kg) NASA ¼ National Aeronautics and Space Administration OD ¼ off-design OEW ¼ operational empty weight (kg) OPR ¼ overall pressure ratio OW ¼ operational weight (kg) PATM ¼ passenger air taxi mission PR ¼ pressure ratio PSR ¼ perfectly stirred reactor SFC ¼ specific fuel consumption (lg/J) TEL ¼ twin-engine light TET ¼ turbine entry temperature UHC ¼ unburned hydrocarbons WGS 84 ¼ world geodetic system dated in 1984

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