c The Society shall not be responsible for statements or opinions advanced in

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1 THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 345 E. 47th St., New York, N.Y c The Society shall not be responsible for statements or opinions advanced in papers or discussion at meetings of the Society or of its Divisions or Sections, or printed in its publications. Discussion is printed only if the paper is published in an ASME Journal. Papers are available from ASME for 15 months after the meeting. Printed in U.S.A. Copyright 1993 by ASME 93-GT-389 CFD STUDY OF NOZZLE CONFIGURATIONS FOR ULTRA HIGH BYPASS ENGINES H. Zimmermann, R. Gumucio, K. Katheder, and A. Jula MTU Motoren- and Turbinen-Union Munchen GmbH Munich, Germany Abstract 1 INTRODUCTION Performance and aerodynamic aspects of ultra-high bypass ratio ducted engines have been investigated with an emphasis on nozzle aerodynamics. The interference with aircraft aerodynamics could not be covered. In recent years computational fluid dynamics (CFD) has become a powerful and more and more generally accepted design tool. For scientific purposes it is sometimes considered of being still too inaccurate, though very often experiments have similar levels of uncertainty. Numerical methods were used for aerodynamic investigations of geometrically different aft end configurations for bypass ratios between 12 and 18, this is the optimum range for long missions which will be important for future civil engine applications. Depending on the geometrical complexity of the type of flow, excellent results and acceptable agreement with test results can be achieved by CFD, see for instance Novak et al, 1992; Dawes, 1992; Jennions and Turner, 1992 and Goyal and Dawes, Results are presented for a wide range of operating conditions and effects on engine performance are discussed. The limitations for higher bypass ratios than 12 to 18 do not come from nozzle aerodynamics but from installation effects. It is shown that using CFD and performance calculations an improved aerodynamic design can be achieved. Based on existing correlations, for thrust and massflow, or using aerodynamic tailoring by CFD and including performance investigations, it is possible to increase the thrust coefficient up to 1%. A number of CFD investigations for engine aft end and nacelle flow, especially those featuring 30 solutions, have been published, see Uenishi et al, 1992; Keith et al, 1992; Brown, 1987 and Norton and Kingsley, Advanced post-processing software tools make it possible to achieve a good representation and to gain a better understanding of the pattern of the complex 3D flow fields. The basic rules for the performance of very high and ultra high bypass ratio turbofans were investigated by Grieb et al, 1986; Zimbrick and Cole- Presented at the International Gas Turbine and Aeroengine Congress and Exposition Cincinnati, Ohio May 24-27, 1993

2 hour, 1990; Skavdahl et al, 1988; Ivey, 1989; Peacock and Sadler, 1989; Eckardt, 1991 and Philpot, But the choice of the cycle depends on thermodynamic and mechanical as well as installation aspects. Further work in this field is necessary before engines with very high and ultra high bypass ratios become available. The aim of the work described in this paper was to optimize the aft end configuration (fan and core nozzle as well as core nozzle cowling) by a comprehensive design process with minimum iterations between aerodynamic (CFD) and performance calculations. As stated in Zimmermann et al, 1992, more than 1.1 PERFORMANCE STUDY Improved thrust specific fuel consumption can be achieved with increased thermal and propulsive efficiencies. The thermal efficiency can be improved by increasing turbine inlet temperature, overall pressure ratio and component efficiency levels. The propulsive efficiency is increased with high bypass ratios corresponding to lower specific thrust. The fan pressure ratio will be reduced. The specific fuel consumption is decreased. 1% can be gained by an optimum design. This is regarded sufficient to justify considerable effort. For an optimized design Dusa et al, 1982, quote The aim of the performance study is the following: a value of 0.5% for the scope of different designs. - the selection of optimum design cycle for NOMENCLATURE Flow subscripts: A - area [m2] a - ambient BPR - bypass ratio [-] e - effective CD - discharge coefficient [-] i - ideal C e- thrust coefficient, eq. (9) [-] max - maximum CF - thrust coefficient, eq. (7) [-] min - minimum C n- thrust coefficient, eq. (10) [-] net - net CV - velocity ratio [-] prop- propulsive C x- ideal thrust coefficient [-] F - gross thrust [kn] Geometrical subscripts: FPR - fan pressure ratio [-] A - afterbody L - distance [m] B - bump Mn - Mach number [-] C - core OPR - overall pressure ratio [-] E - entry P - total pressure [kpa] F - fan p - static pressure [kpa] N - nacelle R - radius [m] 0 - free stream RIT - HPT rotor inlet temperature [K] STP - specific thrust parameters [-] Abbreviations: T - total temperature [K] DP aerodynamic design point TSFC- thrust specific fuel consumption[kg/h/dan] HPC high-pressure compressor V - velocity [m/s] HPT high-pressure turbine Vj - nozzle exit velocity ratio [-] IPC intermediate-pressure compressor (fan to core) ISA international standard atmosphere Vo- flight velocity [m/s] LPT low-pressure turbine w - mass flow rate [kg/s] MCL maximum climb y - ratio of specific heats [-] MCR maximum cruise S - core cowl angle [ ] OD outer diameter n - efficiency [% ] SP sizing point f - fan cowl angle [ ] UHB ultra high bypass ratio p - density [kg/m3] VHB very high bypass ratio 2

3 BPR of 12 to 18 to set a basis for the aerodynamic study to assess performance impact on the design modifications to highlight several performance aspects to identify potential TSFC savings Performance and aerodynamic design study together serve to define the rear end of the compound engine and nacelle. Other aspects have been neglected for this optimization study. 1.2 DESIGN AND INSTALLATION ASPECTS In order to optimize the aft end configuration of ultra high bypass ratio engines for a given performance, it is essential to achieve a maximum gross thrust sum rather than good individual nozzle discharge and thrust coefficients. In this context it is essential to optimize the shape of the core cowl, which has a strong effect on the core thrust (Zimmermann et al, 1992). The nacelle flow (though very important for the magnitude of the nacelle drag) has only a slight influence on the coefficients of the fan nozzle. Accordingly there is no necessity to calculate the full nacelle flow field. The basic configuration for this study is shown in Fig. 1. There, the parameters are indicated, which describe the geometry of an axisymmetric aft end. 2 FLOW FIELD COMPUTATION 2.1 THE CODE The flow field within and around the nozzles was computed with a commercially available CFD code. The numerical method of the "Task Flow" code is described in Raw et al, Features of the code relevant to this investigation are: - strongly conservative - Reynolds averaged Navier Stokes equations, computing turbulent eddy viscosity with the standard k-e model - range from subsonic to supersonic velocities for incompressible and compressible flow - to account for wall friction, the logarithmic law of the wall is available - a high degree of numerical robustness is guaranteed by use of a fully coupled linear solver, accelerated by an additive correction multigrid scheme and a block correction scheme - accuracy is high due to a fully implicit, co-located finite volume method with a flux element-based discretization of geometry and the availability of a second-order discretization scheme. 2.2 THE GRID The computational domain includes the exhaust ducts upstream of the fan and core nozzle exit, the rear part of the nacelle downstream of the maximum diameter, the core cowling, a plug and a sufficiently large ambient. Local obstructions such as pylon and struts are neglected in this axisymmetric grid. The single-block grid is of H-type. Walls are closely described by boundary-fitted coordinates. Near surface grid refinement is used. A flexible grid block structure was developed to cover the different geometries with 224 x 109 nodes. Fig. 1 Geometrical Parameters 2.3 BOUNDARY CONDITIONS A weight study is included, noise and installation aspects are excluded. The design cycle was chosen taking into account some of these criteria. Core, fan and ambient flows are simulated for every bypass ratio at a performance design point corresponding to a flight Mach number of 0.82 at 01

4 35,000 ft. General assumptions for the entire flow field are a constant average viscosity and a constant specific heat capacity. Correct temperatures are specified for the nonswirling, turbulent air and gas streams. is filled by a uniform velocity profile with the ideal velocity derived from the nozzle pressure ratio. This nozzle pressure ratio is formed from the ideal upstream total pressure and the averaged local static pressure in the nozzle exit cross section. Zero flux of mass, momentum and energy is specified at solid wall boundaries such as inner and outer nacelle walls. The logarithmic law of the wall is activated. There are three inflow boundaries: - the upstream far-field boundary, where the flight velocity and a turbulence level of 1% are specified - the fan exhaust duct entry, where the static entry temperature and a turbulence level of 7% are specified Fig. 2 Definition of Subscripts - the low-pressure turbine exhaust, where the total pressure and temperature are specified and a turbulence level of 7% is assumed. In all three cases, homogeneous entry conditions are assumed. In contrast, the conventional discharge coefficient CD a is based on an ideal mass flow or velocity resulting from P/p a, a pressure ratio with the ambient pressure as back pressure CDa = f (P/Pa) (2) The static pressure at the upstream and downstream far-field boundaries is specified, whereas the downstream temperature is extrapolated from nearby cells. The downstream velocity is not specified. A slipline condition is activated for the core jet axis, assuming zero shear and conservation of tangential momentum. It is defined by eq. 2. Physically it is a fictitious value, i.e. it is not consistent with the conditions actually existing in the nozzle. Care has to be taken to avoid errors in using CD a, especially for off-design conditions, where other relationships between P a and nozzle back pressure occur. 3.2 THRUST COEFFICIENTS 3 DEFINITIONS 3.1 DISCHARGE COEFFICIENTS effective mass flow rate w eae CD= = -- = -- (1) ideal mass flow rate wi Ai CD is the ratio of the numerically integrated or measured mass flow rate to that of ideal one-dimensional isentropic flow expansion. It is used to size the nozzle exit cross section for a given mass flow rate. As can be seen from Fig. 2 the effective area For the aerodynamic study, gross thrust only was considered. For the evaluation of thrust generation the ideal thrust coefficient actual thrust F Cx = (3) ideal thrust Fi can be used. There the actual thrust is measured or derived from the axial component of the computed exit velocity in the axially projected exit cross section of the nozzle and the static pressure distribution. F = f Vdw + f (P - Pa) da (4) 4

5 or including correction terms F = CV Cpa -y fp Mn 2 da + f (p -p a )da (5) These correlations are valid for subsonic and sonic nozzle flow. The ideal thrust Fi is the product of ideal flow rate computed from nozzle pressure ratio, and the velocity due to isentropic expansion to ambient pressure P a Fi = wi Vi (6) Depending on the focus of evaluation and to facilitate comparison of high bypass turbofan concepts with regard to a given fan nozzle size, two simple thrust coefficients are proposed. and FF + FA + FC + FN C e = (9) wif ViF FF + FA +FC + FN Cn = (10) 11 RN 2 PF For convenience, the ideal thrust is often based on the measured or calculated real flow rate. Hence, it is possible to obtain the thrust coefficient: CF = actual thrust F ---- _ -- CD a ideal thrust Fi (7) The coefficient is very useful for iterations with the results of the nacelle aerodynamics. In order to assess the efficiency of the whole aft end configuration both coefficients are required. 4. PERFORMANCE INVESTIGATIONS f V dw +f(p - Pa)dA CF= CDa wi Vi or Cx = CF CD a (8) CF over P c/pa is a unique curve for a specific configuration and its value is above unity if in eq. 7, p exceeds P a ; since the numerator is greater than the product of wi and Vi and CD a is always less than one. In the extreme case of w --> 0, CF approaches infinity. Equations 5 and 7 would be appropriate for performance calculations if the nozzle exit static pressure is known from experiment. The load on the outer cone of the core nozzle can be determined from the static pressure distribution on that part. In model tests, usually the sum of the thrusts of both nozzles together with the cone load is measured. It is convenient but not correct to subtract a core thrust representative for a single free nozzle to obtain the fan thrust coefficient. As an improvement it is proposed, that the core load be estimated from numerical calculations. 4.1 SEC STUDY New propulsion concepts such as geared propfans allow a further increase in propulsive efficiency by increasing the bypass ratio and reducing the optimum fan pressure ratio relative to the present in service civil conventional turbofans and also relative to future ungeared engines with very high bypass ratio. Fig. 3 shows the effect of the bypass ratio on installed thrust specific fuel consumption at the Relative TSFC (%) = = Sizing point: MCL, , Mn 0.82, ISA, RIT 1600K, OPR c 40 0 Real Instalretl Engine Referent w.o. 0MAes point Enect ot: I Component EIIic IenGES 2O - Duct Losses Tur_h` Inlet Losses En91nt3 Nonle Thrust Coefficients 30 Iii Installation Drags vne Engines 4O Itleal Cycle UW-Engines _ Bypass Ratio BPR Fig. 3 Thrust Specific Fuel Consumption vs. Bypass Ratio 5

6 sizing point (no customer bleeds, no power extraction) for turbofans and very high bypass ratio engines (BPR < 12) with a pressure ratio of 1.55 to 1.85 in comparison to ultra high bypass ratio engines (BPR > 12), with geared variable pitch fan and a fan pressure ratio of 1.15 to The thermodynamic cycle is based on a core technology of the late nineties, with a HPT rotor inlet temperature of approx K and an overall pressure ratio of approx. 40. The improvement in TSFC attainable from the greater bypass ratio of UHB engines is partly offset by increasing installation losses. The maximum improvement in TSFC (BPR = 20) is on the order of 16% compared with that of a conventional turbofan (BPR = 7), and on the order of 10% compared with that of a VHB engine (BPR = 10). program together with an aircraft company. The turbofan and VHB engine require a thrust reverser with a longer bypass duct and more stages in the LPT, whereas the UHB engine includes a gearbox and fan pitch mechanism. The thrust reverser will be made with variable pitch fan. To evaluate the specific fuel weight per engine, two flight missions were chosen for comparison, namely an estimated short range of about 500 nm and a long range from about 7000 nm. Referred to a long-range mission with optimum bypass ratio of about 15 at the sizing point, the specific total weight (propulsion weight) per engine shows an improvement of about 10% compared with a conventional turbofan and 7% compared with a VHB engine. For short-range missions, the conventionat.turbofan is obviously well established. 4.2 WEIGHT STUDY 4.3 CHOICE OF THE ENGINE FAMILY The substantial improvements in TSFC offered by UHB engines in comparison with conventional turbofans and VHB engines must be balanced against the propulsion weight. It is evident that fan diameter and hence nacelle diameter increase the engine weight. Fig. 4 shows an estimated specific weight per engine versus bypass ratio. A final evaluation should be made by an appropriate aircraft sizing and mission Specific Weight per Engine, Weight / F net. sp (Ibm/Ib) 12 Turbofan & UHS-Engines 11 VHE-Engines 10 Long Range Mission Spec. Fuel Weight per Engine --- Spec. Engine Weight (Estimated) 5Spec. Total (Fuel Engine) Weight per Engine 4 3 Short Range Mission 14 = o Bypass Ratio BPR, SP Fig. 4 Specific Weight per Engine vs. Bypass Ratio With the aim of optimizing the nozzle aerodynamics of aft end configurations, the cycle performance was considered for a family of engines based on the results plotted in Fig. 3 and 4 and with sizing point defined at max. climb, altitude 35,000 feet Mach number 0.82, ISA + 10 C and the same partly installed thrust including ram recovery, customer bleeds and power extraction (no fan- and core-cowl drags). The thermal efficiency was based on more or less the same core technology. The HPC pressure ratio was defined at 10.5 and the average fan face Mach number at 0.68, combined with a hub/tip ratio of 0.4, considering a single rotating fan configuration. By counter rotating fan configurations it is possible to increase the average fan face Mach number to 0.75 and to reduce the hub/tip ratio to In order to study and optimize the aerodynamics of the core and bypass nozzle and the rear end of the engine, 6 different engine cycles have been selected as shown, with some results, in Table 1. Configurations 1 to 3 assume the same thermal efficiency and configurations 4 to 6 the same propulsive efficiency. Configurations 2 and 5 are the identical and the basis for this study.

7 INumber of configuration I I I I ^Bypass ratio X I 1Overall I constant I constant pressure I ratioi I I I I I (Turbine inlet] constant base +160 (temperature I I I [AK]I I lcore size I constant ±2% variable ^AFan diameters-250 base constant [mm]i I ISpec.thrust parameter I 11'an OD pressure I I I I ratio I I I Table 1: Investigated configurations for the sizing point I Relative TSFC (%) Velocity Ratio Vi BPR = V^ / TSFC Aero Design Point: , Mn = 0.82, ISA Fan OD Pressure Ratio FPR Fig. 5 Relative TSFC & Velocity Ratio vs. Fan OD Pressure Ratio Fig. 6 shows the relative TSFC in the bypass ratio range between 12 and 18 in the sizing point. For the given core technology (constant RIT and more or less the same core size), an improvement in the TSFC calls for an increase in the fan diameter i.e. an increase in the propulsor weight. In so doing, it is necessary to bear in mind the airframe installation and aircraft weight. Referring to Fig. 3 and 4, it should be noted that the improvement in the TSFC shows a minimum with a BPR of 20. Considering the total weight for a long-range mission, the minimum is achieved with a BPR of about TRADE-OFF STUDY RELATIVE TO THE SIZING POINT An improvement in TSFC based on increased bypass ratio requires that the fan pressure ratio be optimized at the aerodynamic design point at cruise ISA day. At this condition the efficiencies of the turbocomponents need to be optimised as well. Fig. 5 shows the variation in FPR or velocity ratio Vj (ideal bypass flow velocity/core flow velocity) related to the basic configuration. It is shown that at the minimum TSFC, Vj is typically approximately above 0.8. The required thrust at this point (configurations 2 and 5 with BPR = 16.8) is 79% of the maximum thrust at max. climb. Relative Spec. Fuel Consumption (%) Sizing Point: MCI, , Mn = 0.82, ISA+10 C, F,RIT-eons= Fan Diam. = cons= \ FPR increase FPR = cons= 4 uninst. = cons= 52CtEdPOIRL RIT increase Core Size increase \\\ 215 -'_^ Fan Diam. increase ^3.^ -2 ti I Bypass Ratio BPR Fig. 6 Relative TSFC vs. Bypass Ratio 7

8 Allowing for the total weight for a long-range mission, an improvement in the TSFC for the given fan diameter requires a decrease in core size combined with an increase in the HPT rotor inlet temperature. Optimum would be a BPR of about 16 (Fig. 4). Fig. 7 shows the ratio between core net thrust and total net thrust (net thrust equals gross thrust minus ram drag) in the BPR range from 12 to 18. It is about 14% to 17% in configurations I to 6. Fig. 8 shows the ideal propulsive efficiency and the fan OD pressure ratio versus the specific thrust. Similar to Fig. 6 it is shown that it would be beneficial to go for higher fan diameter but weight effects and installation aspects develop detrimentally. 5 RESULTS OF THE AERODYNAMIC STUDY Relative Core Net Thrust (%) 18 Three investigations were carried out as computational experiments: - variation of bypass ratio from 12 to 18 - modification of the aft end configuration for BPR several aspects of exhaust system aerodynamics The aim was to resolve design and aerodynamic optimisation by computational fluid dynamics. I t Bypass Ratio BPR Fig. 7 Relative Core Net Thrust vs. Bypass Ratio For the constant fan diameter line in Fig. 6, the fan fan pressure ratio and thus the specific thrust and propulsive efficiency are constant. The bypass ratio depends only on the variable core size and hence on the varying thermal efficiency (RIT variable). Ideal Propulsive Efficiency (%) Fan OD Pressure Ratio FPR Sizing Point: MCL, , Mn = 0.82, ISA.10 C /f 3 // 2 j ^^ FPR ^^ 2 ^ 1 Fan Diam. increase Specific Thrust Parameter STP = Fnot uninst / w / V0 Fig. 8 Propulsive Efficiency & FPR vs. Specific Thrust selected point VARIATION OF BYPASS RATIO FROM 12 TO 18 Fig. 9 shows the Mach-number distributions for 4 aft end configurations with a BPR of 12, 15 and 18; the equivalent coefficients, which were obtained by integration of the velocity and static pressure profiles, are given in Table 2. It can be seen that the equivalent data for the 4 configurations is very similar, this is because they were designed to the same principle. The advance in core technology now allows a short design with low losses for the nozzles and cowlings. Configuration 1 in Fig. 9 shows an over-expansion up to Mn = 1.1, this is because of a higher fan nozzle pressure ratio than for the other configurations in Table 2. Whereas the C n coefficient is nearly constant, the C e coefficient increases with fan pressure ratio PF/P a : in the scope of the relevant pressure ratios more thrust can be gained from the pressure term than from the momentum of the velocity profile. Unfortunately, it can be seen from the performance study that configuration 1, with the high fan pressure ratio, has the worst TSFC and must therefore be discarded. 8 7

9 CONFIGURATION a BPR CORE NOZZLE Configuration 1 CD a CD C x CF PC/P a PC/Pmin FAN NOZZLE Configuration 2 CD a CD C x CF PF/P a PF/Pmin Core Cowl [kn] Drag Fan Cowl [kn] Drag C n C e LA/RN Configuration 3 Table 2: Results of CFD Study for Various Bypass Ratios 5.2 MODIFICATIONS OF THE AFT-END CONFIGURATION FOR BPR 12.2 Configuration 4a Fig. 9 Mach-Number Field for BPR 12, 15 and 18 A family of aft end configurations for configuration No. 4 was investigated numerically. The geometries are shown in Fig. 1 and Table 3. The relevant pressure distributions are plotted in Fig. 10. They look very similar since they were designed to the same criteria, i.e. - fan nozzle outer radius constant - the cone angle of the core nozzle cowl stepwise reduced (no flow separation) - core and fan nozzle designed to achieve good CD and C x values 9

10 CONFIGURATION 4a 4b 4c 4d Core cowling ' 11,plug CORE NOZZLE Configuration 4a 8=11 CD a CD C X CF PC/P a PC/Pmin FAN NOZZLE - Configuration 4b 6=14 CD a CD C x CF PF/P a PF/Pmin Configuration 4c 6=6 Core Cowl [kn] Drag Fan Cowl [kn] Drag C n C e LA/RN Table 3: Variation of the Aft End Configuration for BPR 12.2 Configuration 4d 6=11 0 & Plug Fig. 10 Pressure Fields for Different Core Cowling Angles, BPR 12 From Table 3 it can be seen that the C n value for all 4 configurations is almost the same. Consequently the shortest configuration (4d) with the lowest weight is to be preferred, but as well other requirements like installation aspects can easily be covered. Surprising is the small difference in most of the coefficients in Table 3; this was achieved by an optimum design in all cases with good CD values (low losses) and no flow separation. The thrust comes from either the momentum of the velocity profiles, or the nozzle pressure terms or the pressure load on the core cowl, and the three loads balance each other out (i.e. a higher core cowl load is outweighted by lower nozzle thrust). 10

11 A further result of this investigation is that though the overall pressure ratio of the fan nozzle PF/P a is above critical, the real fan nozzle pressure ratio PF/pF is subcritical and consequently the Mach numbers in the fan nozzle are subsonic. This applies for all configurations considered. 5.3 SEVERAL ASPECTS OF EXHAUST SYSTEM AERODYNAMICS (BPR 12.2) Many configurations were investigated for the preparation of this paper but only the optimized configurations are described. As an exception, the Mach-number distribution for a very short and flattened nacelle is shown in Fig. 11. It has the same shape underneath the fan nozzle as in Fig. 10, where the nacelle is axially extended. The short nacelle leads to a design where the bump is outside the fan nozzle, which outweights the thrust balance between the fan nozzle and core cowling. In Table 4 the C n and C e values for configuration 4e are 2 percentage points lower than those for the configuration 4f, which has the same overall thrust as the configurations of Table 3 with a steeper nacelle angle. CONFIGURATION 4d 4e 4f Nacelle normal short long CORE NOZZLE Configuration 4d c=16 & Plug CD a CD C x CF PC/P a PC/Pmin FAN NOZZLE Configuration 4e w =5.5 ShOrt nacelle CD a CD Cx CF PF/P a PF/Pmin Configuration 4f p=5.5 Long nacelle Fig. 11 Mach-Number Fields for Different Nacelle Concepts Core Cowl [kn] Drag Fan Cowl [kn] Drag C n C e LA/RN LB/RB Table 4: Comparison of Several Configurations with Core Plug or Bump (see Fig. 1) 11

12 It can be concluded that if nacelle, nozzles and core cowling are designed to basic rules of low losses, minimum base area and no flow separation, the differences in thrust for a given performance will be less than 1%. In Fig. 12 the thrust coefficients CF and C x of the core nozzle of configuration 4b are plotted against P c/pa is For the latter case, when related to the net thrust the percentage is approximately 86 (see Fig. 7). of gross thrust goo -1 90,1 % 1 Fan nozzle thrust 2 Core nozzle thrust 3 Core cowling pressure thrust 4 Nacelle pressure drag 5 Nacelle friction drag 6 Core cowling friction drag Cr, Cx 1 CF +F 10,7% F Thrust Fig. coefficients 13 Thrust Balance Configuration 4a 0.90 CX BPR = 12.2 of gross thrust 100 ^ 92,8 % r T --- i 1 I PC/Pa 1.0 1,2 1, Core nozzle pressure ratio 1 Fan nozzle thrust 2 Core nozzle thrust 3 Core cowling pressure thrust 4 Nacelle pressure drag 5 Nacelle friction drag 6 Core cowling friction drag Fig. 12 Core Nozzle Thrust Coefficients tf 0 8% As eq. 8 shows CF = Cx/CDa -F Fig. 14 Thrust Balance Configuration 3 BPR = 18.2 Since C x and CD a both have been derived by integration from the same velocity and pressure profile, they are similarly dependent on the nozzle pressure ratio. Consequently CF is a unique curve and greater than unity, as explained in section 3.2. Cx very much depends on the fan flow field and is additionally a function of PF/PC Fig. 13 and 14 show a break down of thrusts, drags and friction forces. It can be seen that the friction forces are negligible and that the force on the core cowling is positive (thrust) for all configurations. The drag on the rear of the fan cowling is relatively small. For configuration 4a with BPR = 12.2 the fan thrust dominates with 90.1%, for configuration 3 with BPR = 18.2 the fan thrust percentage 6. CONCLUSIONS - Performance and weight trade off studies show an optimum in TSFC between BPR 12 and 18 for longrange missions. - The fan pressure ratio has to be adjusted, the velocity ratio Vj (fan to core) is of the order 0.8 at the aerodynamic design point. - In the range of BPR 12 to 18 an improved TSFC can be achieved by lower RIT and a higher fan diameter but limitations come from weight and installation aspects. 12

13 If the whole aft end configuration is designed by aerodynamic optimization, the difference between alternative designs is less than 1% in TSFC. The aft end can therefore be optimized based on installation and weight aspects. The progress in core technology allows a short, lightweight design without drag penalties. Especially short designs are feasible with a plug in the core nozzle. Because of afterexpansion effects behind the fan nozzle, the fan nozzle exit velocity was subsonic for all configurations considered. REFERENCES Brown, J.J. "Navier-Stokes Analysis of a Very-High Bypass Ratio Turbofan Engine in Reverse Thrust" AIAA Paper , 1987 Dawes, W.N., "The Extension of a Solution-Adaptive 3D Navier-Stokes Solver Towards Geometries of Arbitrary Complexity" ASME Paper 92-GT-363, 1992 Dusa, D., Lahti, D.J., Berry, D. "Investigation of Subsonic Nacelle Performance Improvement Concept" AIAA Paper , 1982 Eckardt, D. "Future Engine Design Trade Offs" X. ISABE, Nottingham, U K., Sept.1-6, 1991 Goyal, R.K., Dawes, W.N. "A Comparison of the Measured and Predicted Flowfield in a Modern Fan-Bypass Configuration" ASME Paper 92-GT-298, 1992 Grieb, H., Eckardt, D. "Turbofan and Propfan as Basis for Future Economic Propulsion Concepts" AIAA Paper , 1986 Ivey, M.S. "The Impact of Bypass Ratio on the Performance of Future Civil Aero Engines" Aerotech 89, Seminar 15, Birmingham, 1989 Jennions, I.K., Turner, M.G. "Three-Dimensional Navier-Stockes Computations of Transonic Fan Flow Using an Explicit Flow Solver and an Implicit n - e Solver" ASME Paper 92-GT-309, 1992 Keith, B.D., Uenishi, K., Dietrich, D.A. "CFD-Based 3D Turbofan Exhaust Nozzle-Analysis System" AIAA Paper , 1991 Norton, R.J.G., Kingsley, J.P. "Prediction of a High Bypass Ratio Engine Exhaust Nozzle Flow Field" AIAA Paper , 1992 Novak, 0., Schafer, 0., Schonung, B., Patzold, H. "Use of Advanced CFD Codes in the Turbomachinery Design Process" ASME Paper 92-GT-324, 1992 Peacock, N.J., Sadler, J.H.R., "Advanced Propulsion Systems for Large Subsonic Transports" AIAA Paper , 1989 Philpot, M.G. "Practical Considerations in Designing the Engine Cycle" AGARD-LS-183, pp. 2-1 to 2-24, 1993 Raw, M.J., Galpin, P.F., Hutchinson, B.R., "A Colocated Finite Volume Method for Solving the Navier-Stokes Equations for Incompressible and Compressible Flows in Turbomachinery: Results and Applications" Annual General Meeting of the Canadian Aeronautics and Space Institute, 36th, Ottawa,

14 Skavdahl, H., Zimbrick, R.A., Colehour, J.L., Sallee, G.P. "Very High Bypass Ratio Engines for Commercial Transport Propulsion" ICAS , pp , 1988 Uenishi, K., Pearson, M.S., Lehnig, T.R., Leon, R.M. "Computational Fluid Dynamics Based Three-Dimensional Turbofan Inlet/Fan Cowl Analysis System" Journ. of Propulsion and Power, Vol. 8, No. 1, Jan.-Feb Zimbrick, R.A., Colehour, J.L., "Investigation of Very High Bypass Ratio Engines for Subsonic Transports" J. Propulsion, Vol. 6, No. 4, 1990, pp Zimmermann, H., Katheder, K., Jula, A. "A Numerical Investigation into the Nozzle Flow of High By-Pass Turbofans" ASME Paper 92-GT-10 (1992) 14