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ISABE-2015-20090

OPTIMIZATION OF AN ENGINE WITH A GEAR DRIVEN COUNTER ROTATING FAN PART II: CYCLE SELECTION AND PERFORMANCE

Tom Otten, Timea Lengyel-Kampmann, Richard Becker, Stanislaus Reitenbach

German Aerospace Center (DLR) Institute of Propulsion Technology

Abstract

Aircraft fuel efficiency becomes more and more important because of increasing fuel prices and environmental concerns. Engine efficiency contributes substantially to this, but further improvements are very demanding due to the already mature technology. To achieve further enhancements, the introduction of new engine concepts is currently in discussion. Most of these concepts focus on improving propulsive efficiency.

In this two-part study, the concept of the counter rotating fan (CR-fan) is assessed. In the first part of this study [1], the fan concept is analysed considering aerodynamical aspects. The focus of the present part of the study is to assess the CR-fan concept on engine level.

To assess the results of the aerodynamic optimization on an overall engine level, a cycle design study is performed. For a civil long range application, the aircraft characteristics are modelled and engine requirements are derived. Then a cycle is developed, taking both the constraints from the thermodynamic limitations and the thrust requirements from the flight mission into account.

Besides the thermodynamic modelling, weight and drag analyses for the engine are performed.

The fan aerodynamics and weight is directly taken from the first part of this study [1]. The technology assumptions for the other engine components are taken from literature. The close link between fan design and performance study allows for a detailed description of the propulsor within the thermodynamic model.

As a result the optimal aerodynamic fan design is identified by finding the best compromise between cycle demands, fan efficiency, weight and drag in order to minimize fuel consumption.

Finally, an engine cycle for a CR-fan is defined taking into account the gearbox influenced off-design behaviour of the fan rotors.

Nomenclature

CR-fan Counter Rotating Fan BPR Bypass Ratio D Drag EoF End of Field FPR Fan Pressure Ratio HPC High Pressure Compressor HPT High Pressure Turbine i0 Base Gear Ratio L/D Lift to Drag Ratio LPC Low Pressure Compressor LPshaft Low Pressure Shaft LPT Low Pressure Turbine MTOW Maximum Take-off Weight OEW Operating Empty Weight OPR Overall Pressure Ratio PDG Planetary Differential Gear P Power RN Reynolds Number TET Turbine Entry Temperature TOC Top of Climb TSFC Thrust-specific Fuel Consumption u Circumferential Speed 𝑣𝑣𝑖𝑖𝑖𝑖 Ideal Speed W Weight XN Rotational Speed XM Mach Number α Flow Angle π Pressure Ratio 𝜂𝜂𝑖𝑖𝑖𝑖 Isentropic Efficiency Subscripts

Byp Bypass Core Core MC Mid Cruise Mod Modified Rel Relative Ref Reference

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Introduction

Counter rotating fans (CR-fans) have the potential to improve the overall efficiency of today’s aero engines [2]. By splitting the pressure rise into two rotor rows, low blade numbers per row and high axial rotor inlet Mach numbers can be achieved, enabling lower fan diameters compared to direct driven or geared turbofans. As the second rotor removes most of the swirl, no stator is required for this concept. The absence of stator losses and the low stage loadings enable very high fan efficiencies and high bypass ratios which correspond to high propulsive efficiencies.

A counter rotating fan can be driven in various ways: Besides counter rotating turbines, the planetary differential gear (PDG) was found to be a suitable solution to drive a CR-fan. Both options are also in discussion to drive open rotors [3].

Moreover, following the introduction of the geared turbofan engines (GTF) by Pratt and Whitney, a planetary gear drive has gained acceptance as an engine module [4].

In the first part of the present publication [1], the aerodynamic performance of the PDG driven counter rotating fan is analysed for a large range of axial fan entry Mach numbers (S2.XM) and fan pressure ratios (FPR) using a generalized 3D-optimization utilizing a RANS-CFD-Solver. Furthermore, the weight of the fan, including blades, disks and fan case is evaluated for each design in the parameter range.

The focus of this part of the present study is to assess the CR-fan on engine level, including the gearbox effects as well as its’ weight and drag.

Methodology

To evaluate the CR-fan on an overall level, an engine cycle analysis is performed. Furthermore, a method to derive the weight and drag of the overall engine is applied. Then a flight mission analysis is conducted to define a figure of merit that includes the engine performance, its weight, drag and efficiency.

Engine Performance The evaluation of the CR-fan on engine level is performed within DLR’s in-house performance code GTlab-Performance [5].

The engine cycle is designed as a 2-spool unmixed turbofan [Figure 1]. The core engine components are modelled using standard component maps. The high-pressure turbine (HPT) is assumed as a two stage cooled turbine while the low pressure turbine (LPT) is uncooled.

The cooling air is taken from two stations within the high pressure compressor (HPC). The first HPT stage cooling air is taken from the HPC exit,

while the second stage cooling air is taken from a bleed port at 80% of HPC enthalpy rise.

Figure 1: Thermodynamic cycle schematic of the CR-fan engine

The efficiency of the compressors and the LPT is modelled as a function of corrected mass flow as described in [6], while for the HPT, the effect of cooling air on efficiency is included as well [6].

The bypass ratio (BPR) is defined by matching the nozzle speed ratio to the ideal nozzle speed ratio as suggested in [7].

𝑣𝑣𝑖𝑖𝑖𝑖,𝐵𝐵𝐵𝐵𝐵𝐵

𝑣𝑣𝑖𝑖𝑖𝑖,𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶= 𝜂𝜂𝑖𝑖𝑖𝑖,𝐿𝐿𝐿𝐿𝐿𝐿 ∗ 𝜂𝜂𝑖𝑖𝑖𝑖,𝐹𝐹𝐹𝐹𝐹𝐹 (1)

The fan mass flow is adapted to match the

required thrust. A constant pressure ratio split between low pressure compressor (LPC) and HPC is assumed.

The cycle is designed for cruise condition, as this is the most important point in terms of fuel consumption, especially for a long range design. Besides the design point, relevant off-design points such as End of Field (EoF) and Top of Climb (TOC) are taken into account in order to verify maximum thermal and mechanical loads.

In the following, the FPR will be a free parameter. Especially for low FPR, a variable area bypass nozzle will become necessary to ensure save surge margins at critical off-design points like Take-off/EoF condition. The nozzle area is adapted to match a given distance from surge line in the fan performance map in order to ensure stable operation.

Fan and Gearbox modelling The CR-fan description for the engine performance simulation must represent the design- as well as the off-design behavior of both rotors and is closely linked to the gearbox description.

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Figure 2: Counter Rotating Fan

The CR-fan [Figure 2] consists of two fan stages, rotating in opposite direction. In this paper, it is assumed, that both rotors are driven by a PDG as discussed in [1].

This configuration inevitably leads to an exit swirl after Rotor 2. This swirl leads to a decreased thrust output compared to a purely axial flow, thus it has to be considered as a loss. A modified efficiency 𝜂𝜂𝑖𝑖𝑖𝑖,𝑚𝑚𝐶𝐶𝑖𝑖 ([1], [8]) was used to describe the losses of the exit swirl.

Figure 3: Module representation of the CR-Fan within GTlab-Performance

There are several ways to describe the CR-fan within engine performance calculations [8]. In this approach, both rotors are described as one module that incorporates the gear [Figure 3]. This option offers sufficient flexibility without the need to model the interaction between both rotors in detail.

The design point performance of the CR-fan is directly taken from response surfaces (Equation 2) created in [1].

𝜂𝜂𝑖𝑖𝑖𝑖,𝑚𝑚𝐶𝐶𝑖𝑖𝑖𝑖0

𝑋𝑋𝑋𝑋𝑅𝑅1𝑋𝑋𝑋𝑋𝑅𝑅2

𝑋𝑋𝑋𝑋𝑅𝑅1,𝑅𝑅𝐶𝐶𝑅𝑅⎭⎪⎬

⎪⎫

= 𝑓𝑓(𝐹𝐹𝐹𝐹𝐹𝐹, 𝑆𝑆2.𝑋𝑋𝑋𝑋) (2)

The response surface for the modified efficiency is shown in Figure 4.

Figure 4: Modified efficiency response surface. Difference to reference value ∆η = η-ηref in %-points is shown. (from [1])

Besides the modified efficiency, the base gear ratio i0 and the rotational speed ratio between both rotors is used to calculate the speed ratio between LPT and the rotors. In [1] the fans are calculated for a constant diameter of 3 meter, while in this study, the fan diameter is supposed to vary, so the rotational speed 𝑋𝑋𝑋𝑋𝑅𝑅1,𝑅𝑅𝐶𝐶𝑅𝑅 has to be corrected to ensure the right tip speed.

The off-design behavior of the counter rotating fan is described by using a scaled performance map that includes the characteristic of the PDG [8]. Herein the fan performance data is stored in tables depending on 𝑋𝑋𝑋𝑋𝑅𝑅1 and an auxiliary coordinate β as suggested for conventional compressors [20]. Besides the standard map tables like efficiency, pressure ratio and mass flow, the CR-fan map has additional tables i.e. for the Rotor 2 speed.

Engine Weight Estimation The evaluation of the best engine setup is not solely defined by engine efficiency. The engine weight contributes significantly to the efficiency of the overall aircraft system. The determination of the total engine weight is inherently difficult in this stage of development, as only brief descriptions of its components are available.

The weight and drag estimation of this study is based on the results of a conceptual flow path design process. Based on thermodynamic cycle parameter

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and additional inputs like Mach numbers and hub-to-tip ratios at all stations, a preliminary annulus design is derived. For the core engine, these values were fixed during the parameter study.

To derive the compressor and turbine stage number, average stage loadings (Table 1) are assumed based on [6],[9].

Component Work coefficient

𝚿𝚿� = 𝟐𝟐∗∆𝒉𝒉𝒖𝒖𝟐𝟐

LPC 0.5 HPC 0.5 LPT 1.7

Table 1: Stage loadings

The resulting component annulus approximations are assembled to an overall bare engine flow path which provides the input parameters for the subsequent components weight and nacelle geometry calculations.

In this study, the approach of Sagerser [10] is used to determine the weight of engine core components. It’s applicability to modern engines was evaluated i.e. in [11]. The component weights derived by Sagerser’s method are corrected to match internal reference data as well as public engine weights [12]. The determination of the engine nacelle weight is performed by a simple correlation developed in [13]. The fan weight is directly taken from the first part [1] and scaled to the required diameter, using the gradients from [10]. The gearbox weight was estimated using a correlation from [14].

Engine drag The drag of the engine nacelle is determined by means of the component build up methodology proposed by [15]. The zero-lift drag coefficient is calculated using equation (3)

𝐶𝐶𝑖𝑖,0 =𝐶𝐶𝑅𝑅 ∗ 𝐹𝐹𝐹𝐹 ∗ 𝐼𝐼𝐹𝐹 ∗ 𝑆𝑆𝑤𝑤𝐶𝐶𝑤𝑤

𝑆𝑆𝐶𝐶𝐶𝐶𝑅𝑅 (3)

The interference factor IF was fixed to 1.3 in

this study. The form factor is defined by equation (4)

𝐹𝐹𝐹𝐹 = 1 + �0.35𝑑𝑑𝐹𝐹𝐹𝐹𝑛𝑛𝑙𝑙𝐹𝐹𝐹𝐹𝑛𝑛

� (4)

The maximum nacelle diameter dnac, overall

nacelle length lnac and wetted area Swet are derived from the nacelle geometry computations [13].

The flat-plate skin friction coefficient for turbulent flow was determined by equation (5).

𝐶𝐶𝑅𝑅 =0.455

(log10 𝐹𝐹𝑋𝑋)2.58 ∗ (1 + 0.144 ∗ 𝑋𝑋𝑋𝑋2)0.65 (5)

Flight Mission Simulation Typically, the use case for an engine is a RFP (Request for proposal) for a fixed airplane with specific requirements for engine efficiency and weight.

Adapted to a fixed aircraft, a change in engine efficiency, weight or drag would lead to different aircraft operating distances or to a change in aircraft payload. These influences are difficult to balance and thus, the potential of the regarded engine concept cannot be evaluated easily.

In this study another approach is selected, as the efficiency and weight of the engine will vary and must be set in relation to one another. The flight mission analysis poses a possibility to define a suitable figure of merit. Hence, a fixed design mission is defined, while the aircraft characteristic is iteratively adapted to the engine design as shown in Figure 5.

Figure 5: Schematics of the interaction between aircraft and engine in the optimization process

A future wide body airplane is selected as baseline configuration. The operating empty weight (OEW) is set to 108501kg, based on early design estimations of the Boeing 787 [16]. The reference dry engine weight of 5048kg/engine is taken from the same study. It should be noted, that all these assumptions are focused on representing a realistic modern aircraft rather than matching the exact Boeing 787 performance.

The mission calculation is performed backwards. The descend segment is neglected as it has a very small effect on fuel burn for a long range mission. The Breguet equation (Equation 6)

𝐿𝐿 = 𝑣𝑣𝐹𝐹𝐹𝐹𝑖𝑖𝐹𝐹ℎ𝑤𝑤 �𝐿𝐿 𝐷𝐷�

������𝑛𝑛𝐶𝐶𝐶𝐶𝐶𝐶

𝑔𝑔 ∗ 𝑆𝑆𝐹𝐹𝐶𝐶����� ∗ ln𝑊𝑊𝐿𝐿

𝑊𝑊𝐶𝐶𝑅𝑅,𝑖𝑖𝐹𝐹𝑖𝑖𝑤𝑤 (6)

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is applied to calculate the weight at initial cruise 𝑊𝑊𝐶𝐶𝑅𝑅,𝑖𝑖𝐹𝐹𝑖𝑖𝑤𝑤 for a given mission length L of 8400nm.

The landing weight (𝑊𝑊𝐿𝐿) depends on the operating empty weight, the payload and the reserve fuel (Equation 7):

𝑊𝑊𝐿𝐿 = 𝑂𝑂𝑂𝑂𝑊𝑊 + 𝐹𝐹𝑃𝑃𝑃𝑃𝑙𝑙𝑃𝑃𝑃𝑃𝑑𝑑 + 𝐹𝐹𝑅𝑅𝑅𝑅𝑅𝑅𝑅𝑅𝑣𝑣𝑅𝑅𝐹𝐹𝑅𝑅𝑅𝑅𝑙𝑙 (7)

For the mission, a constant payload of 21.34t

is assumed. The OEW includes the engine and nacelle weight. The reference weights are replaced by the calculated weights for the new engine (Equation 8). 𝑂𝑂𝑂𝑂𝑊𝑊 = 𝑂𝑂𝑂𝑂𝑊𝑊𝐶𝐶𝐶𝐶𝑅𝑅 + 𝑆𝑆𝑆𝑆𝐹𝐹 ∗ (𝑊𝑊𝐸𝐸𝐹𝐹𝐹𝐹,𝑅𝑅𝐶𝐶𝑅𝑅 −𝑊𝑊𝐸𝐸𝐹𝐹𝐹𝐹,𝑅𝑅𝐶𝐶𝑅𝑅) (8)

In reference [17] the factor describing the

influence of engine weight on aircraft structural weight (SBF) is assumed in a range between 1.5 and 4. In this study this correlation factor is set to 2.

The reserve fuel amount consists of three fractions:

• 5% of total fuel burn • Fuel burn for 30minutes of flight • Fuel burn for the diversion to an alternate

airfield (200nm distance).

The lift-to-drag ratio �𝐿𝐿 𝐷𝐷�������

𝑛𝑛𝐶𝐶𝐶𝐶𝐶𝐶 is assumed to be

constant during cruise. The base L/D was set to 20.84 [16]. The drag difference Δ𝐷𝐷 between the reference engines and the current engines is then iteratively applied to the L/D by determining the aircraft mass 𝑊𝑊𝑀𝑀𝐶𝐶 at mid cruise using the Breguet equation for half the mission length. The corrected L/D can then be derived by equation 9:

�𝐿𝐿 𝐷𝐷� �𝑛𝑛𝐶𝐶𝐶𝐶𝐶𝐶

= 𝑔𝑔 ∗𝑊𝑊𝑀𝑀𝐶𝐶

𝑔𝑔 ∗ 𝑊𝑊𝑀𝑀𝐶𝐶𝐿𝐿𝐷𝐷�

+ Δ𝐷𝐷

(9)

Furthermore a constant specific fuel burn TSFC is calculated for cruise condition by means of the engine performance simulation.

As a result, the aircraft weight as well as the engine thrust at the beginning of the cruise segment is known. Assuming a climb rate of 650ft/min, the thrust requirement at TOC can be calculated.

For the climb segment, a fixed duration of 28 minutes is assumed. Based on engine performance calculations, an averaged climb thrust and fuel flow is determined, so the aircraft weight difference between cruise and take-off can be estimated, resulting in the definition of the aircraft maximum take-off weight (MTOW).

The EoF thrust requirement is calculated at zero altitude, a flight Mach number of 0.25 and a

temperature delta to ISA condition of +15K using MTOW. As in this condition flap and landing gear are deployed, an adopted L/D of 16.58 (according to [16]) and a climb angle of 8.2° are assumed in order to calculate the EoF thrust.

With this approach, the thrust requirements for the engine cycle are defined iteratively as they depend on the engine. The engine size influences the drag and thus reduces L/D, while the engine weight affects aircraft OEW.

It should be noted that this mission calculation is not intended to represent exact existing aircraft performance, but offers a simple possibility to compare engines on an overall aircraft level at very little time. The assumptions on aircraft and mission performance are summarized in Table 2.

L/D @ Cruise - 20.84

L/D @ EoF - 16.58

Rate of Climb @ TOC ft/min 640

ClimbAngle @ EoF ° 8.2

Reference Area m² 359.5

OEW kg 108501

Payload kg 21337 (224Pax)

Mission Length nm 8400

Holding Time min 30

Dst. To Alternate nm 200

Ref. Engine weight kg 5048

Ref. Nacelle weight kg 1951

Table 2: Aircraft assumptions, taken from [16]

Results

The study is performed in three steps. First, the uninstalled performance is derived, only considering the engine. Then, the mission based performance is evaluated for constant design-point cycle parameters. In the last step, these cycle parameters are altered to match off-design temperature limits.

Uninstalled performance In this setup, the thrust requirements of the engine are fixed and the engine cycle is set to a constant overall pressure ratio (OPR) and turbine entry temperature (TET). The FPR and the axial fan inlet Mach number are varied and a cycle calculation is performed for each combination. When varying the FPR, the LPC and HPC pressure ratios are adapted to ensure a

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constant OPR. The pressure ratio split between both compressors is kept constant.

Figure 6 shows the engine TSFC delta for different fan pressure ratios and axial fan inlet Mach numbers. Minimum TSFC is found at the lowest considered fan pressure ratio and Mach number. Aerodynamic analysis indicated optimal fan efficiencies (Figure 4) around FPR 1.3 and low values for S2.XM. However, a further decrease in FPR leads to higher propulsive efficiency that outperforms the effects of the fan efficiency on the overall performance.

The influence of the axial fan entry Mach number is comparatively small. Due to the higher fan efficiency, the lowest S2.XM show slightly better SFC values.

Figure 6: Cruise TSFC delta as a function of FPR and S2.XM for constant OPR, TET and Thrust

As expected, the engines’ bypass ratio [Figure 7] increases with decreasing fan pressure ratios. The considered range of FPRs corresponds to BPRs between 13 and 30.

Figure 7: Bypass Ratio as a function of FPR and S2.XM for constant OPR, TET and Thrust

To keep the thrust constant, the fan mass flow has to increase with lower FPR. This effect is shown in Figure 8 as the required fan diameter increases with a decrease in FPR. A higher S2.XM enables a reduction in fan diameter. Increasing S2.XM from 0.6 to 0.725, the diameter is reduced by 0.15 - 0.2m.

Figure 8: Fan Diameter as a function of FPR and S2.XM for constant OPR, TET and Thrust

In Figure 9 the derived engine system weight is shown. The engine weight increases strongly with decreasing FPR. The weight increase is mostly driven by the fan system and nacelle weight.

The weight gradient becomes larger at low FPR (<1.25). A higher S2.XM helps to keep the engine weight low as it reduces the fan diameter.

Figure 9: Total Engine weight (incl. Nacelle) as a function of FPR and S2.XM for constant OPR, TET and Thrust

Besides weight, the engine drag is of importance to evaluate the engine concept. In Figure 10 the absolute drag per engine at cruise is shown. Because the drag is modelled as a function of the wetted nacelle area, the drag increases with the fan diameter.

+0.05g/kNs

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Figure 10: Engine Drag at Cruise as a function of FPR and S2.XM for constant OPR, TET and Thrust

Installed performance, constant cycle To compare the weight and drag with the engine efficiency, the results on a flight mission are important.

While the TSFC directly influences the mission fuel burn, it also has an indirect influence as well: The better the fuel consumption the lower the aircraft weight (and thus thrust requirement) will be. The heavier the engine gets, the higher the thrust requirement will be.

By applying the abovementioned methodology, a total mission fuel burn can be calculated that inherits all these effects.

The results for the uninstalled engine in terms of TSFC remain practically constant, as the main cycle parameters are kept unchanged. As the thrust requirement of the engine changes, the size of the engine is varied.

Figure 11: Total Mission Fuel Burn as a function of FPR and S2.XM for constant OPR and TET

The total mission fuel burn is shown in Figure

11 for different S2.XM and FPR. Towards very low fan pressure ratios a large increase in fuel burn is found. In this area, the decreasing fan efficiency and the increase in weight and drag dominates the increase in propulsive efficiency by far.

It can be derived, that the engine cycle that offers the best mission fuel burn is found for a fan pressure ratio of 1.325 and an S2.XM of 0.68.

Figure 12: Thrust requirement at initial cruise

The thrust requirement in this setup varies due to both the fuel and engine weight and its drag. Its trend at initial cruise is presented in Figure 12.

The required thrust increases with decreasing FPR due to the higher mission fuel burn and the higher engine weight and drag. A higher S2.XM decreases the thrust requirement.

The assumption of a constant OPR and constant TET at cruise leads to a different off-design behaviour due to the higher thrust lapse of the low FPR cycles. As the limits for the engine cycle are typically found at off-design conditions, particularly with regard to TOC and EoF, this effect must be taken into account.

Especially for the regarded high-BPR cycles, the EoF condition poses the highest temperatures and mechanical limitations and thus limits the cycle design. In this study, the HPC exit Temperature (T3) and the LPT inlet temperature (T45) are found to be the most important constraints. For the constant engine cycle, the variation of T3 and T45 at EoF is shown in Figure 13. The maximum occurring temperature delta for T3 is 32K and for T45 64K.

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Figure 13: T3 and T45 at End of Field Condition

Adapted Cycle and Thrust The results in the previous chapter are based on constant design (cruise) values for OPR and TET. To ensure an assessment on the same technological constraints, the engine cycle parameters are varied to achieve the same temperature and mechanical limits at off-design operation. The highest temperatures are found at EoF condition while the highest rotational speeds stay below critical limits and are not considered in the following. The engine OPR and TET are adapted until the T3 and T45 limits are met.

For T3, the maximum allowed temperature is set to 970K according to [18]. To enable an uncooled LPT, a maximum LPT inlet temperature of T45=1250K is set [18].

Figure 14: Development of Engine OPR and Cruise TET at cruise for various fan pressure ratios.

As the off-design temperatures at high FPR exceed the allowed limitations, the engine cycle requires lower TET and OPR at Design conditions. For low FPR, the OPR and TET can be increased. The influence of S2.XM on TET and OPR is comparably small.

In Figure 14 the engine OPR and TET is shown over the fan pressure ratio. The OPR varies from 50.4 (at the highest FPR) to 55.7 for the lowest FPR.

In Figure 15 the resulting total mission fuel burn is shown. The optimum FPR and S2.XM remains unaffected by the variation in OPR and TET. Towards higher FPR, the mission fuel burn increases stronger than for the constant cycle. This is plausible, as the lower OPR and TET will reduce the cycle efficiency. The high gradient in mission fuel burn at FPR~1.28 is due to a stage number change in the LPT.

Figure 15: Total Mission Fuel Burn for an adapted cycle.

Sensitivity study on engine weight To cope for uncertainties especially regarding engine weight, the same calculation is performed for an engine weight 10% higher and lower as originally calculated. As expected, the results (Figure 16, Figure 17) show a higher total fuel burn value for the heavier engine whereas for the lighter engine, the total fuel burn is reduced. The optimum fan design in terms of FPR and S2.XM is not shifted significantly, only a minor increase in ideal S2.XM (0.68 => 0.69) and FPR (1.3 => 1.31) can be observed towards higher weight assumptions.

Figure 16: Total Mission Fuel Burn for 10% heavier engine

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Figure 17: Total Mission Fuel Burn for 10% lighter engine

Conclusions and Outlook

A methodology to evaluate a new fan concept on flight mission level was presented and adopted to a counter rotating fan. For this, different fan designs were obtained from an aerodynamic design study [1] and implemented into the engine cycle design process as response surfaces. A weight and drag analysis was performed based on correlation methods and on data taken from the component design. These results were combined in a flight mission calculation which matches the thrust requirement, the engine weight and drag as well as the engine fuel burn iteratively. The resulting mission fuel burn can be used as a figure of merit for regarded engine concept.

It could be shown, that the fan with the best efficiency is not automatically the optimal member for the overall engine. The effects of propulsive efficiency, fan efficiency, weight and drag contribute substantially to the definition of the optimal fan.

The best fan member is largely defined by the fan pressure ratio, which showed highest efficiencies for a FPR of 1.32. The axial fan inlet Mach number has a smaller effect on fuel burn than the FPR but should be taken into account when selecting the ideal fan for an engine concept. The best axial fan inlet Mach number was found at 0.68-0.7.

This study concentrated on the development of the mission based evaluation of the CR-fan. A comparison with a conventional fan will offer the possibility to assess the advantages and drawbacks of the counter rotating concept in detail.

Currently, the engine conceptual design tool GTlab-Sketchpad [19] is in development at DLR and will improve the reliability of the geometry and weight assessments. A direct coupling of the engine performance software GTlab-Performance and the

aerodynamic predesign and design tools will further enhance the flexibility of the analysis.

References

[1] Lengyel-Kampmann T.; Otten, T.; Schmidt, T.; Nicke, E.: Optimization of an Engine with a gear driven counter rotating fan PART I: Fan Performance and Design; ISABE-2015-20091, 22th ISABE Conference Phoenix, Arizona, USA, 2015

[2] Otten, T.; Becker, R. ; Plohr, M.; Doepelheuer, A.; Energy efficient Engine Concepts, NATO CSO AVT Workshop on Energy Efficient Technologies and Concepts of Operation; Lisbon, Portugal, 22-24. October 2012

[3] Dron, S.: Towards ACARE 2020: Innovative Engine Architectures to Achieve the Environmental Goals; 26th International Congress of the Aeronautical Sciences, ICAS, 2008

[4] Riegler, C.; Bichlmaier, C.; The Geared Turbofan Technology - Opportunities, Challenges and Readiness Status, Proceedings of the 1st CEAS European Air and Space Conference, Berlin, 2007

[5] Becker, R.; Wolters, F.; Nauroz, M.; Otten, T.: Development of a Gas Turbine Performance Code and its Application to Preliminary Engine Design, DLRK2011-24485, DGLR, Deutscher Luft- und Raumfahrtkongress Bremen, 2011

[6] Grieb, H.: Projektierung von Turboflugtriebwerken, Birkhauser Verlag, 2004

[7] Kurzke, J.: Aeroengine Design: From State of the art turbofans towards innovative architectures, VKI Lecture Series; von Karman Institute for Fluid Dynamics, 2013

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