Optimization Study of An Intercooled-Recuperated Engine

1 ISABE-2011-1318 ANALYSIS OF AN INTERCOOLED RECUPERATED AERO-ENGINE Xu Lei Chalmers University of Technology Gothenburg 41296, Sweden

Konstantinos Kyprianidis Chalmers University of Technology Gothenburg 41296, Sweden

Tomas Grönstedt Chalmers University of Technology Gothenburg 41296, Sweden

Abstract The multi-disciplinary tool TERA2020 is employed to explore the design space of the Intercooled Recuperated Aero-engine being studied in the NEWAC project. To have a better understanding of the design space, a parametric analysis was conducted. This allows selecting of an engine close to optimal performance. Within the constraints, a 4.61% of fuel saving was observed compared to the baseline engine. Nomenclature A CR T TO TOC W ε η π

area mid cruise temperature take off top of climb mass flow rate effectiveness efficiency pressure ratio

Subscripts m 1 23 3 31 40 41

metal temperature fan inlet low pressure compressor inlet high pressure compressor outlet recuperator inlet combustor outlet high pressure turbine rotor inlet 5 low pressure turbine inlet 132 intercooler inlet

Introduction The Intercooled Recuperated Aeroengine (IRA) has been proposed and studied for long time [1][2][3]. In the NEWAC [4] project the IRA engine studied in the earlier projects [5][6] is further developed to achieve a higher technical readiness level and optimize its performance [7]. The involved components are refined and studied in more detail. Meanwhile, the engine is modeled by the university partners (Chalmers University of Technology, Stuttgart University, Cranfield University and National Technical University of Athens) and system level investigation is carried out to explore the design space. The IRA Engine and design variables Compared to a conventional engine, the NEWAC IRA engine configuration is more complex. Besides the intercooler (IC) and recuperator (REC), this engine also features the introduction of variable guide vanes (VGVs) in the low pressure turbine (LPT), as illustrated in Figure 1. These new components and functionality offer much more design freedom, but the search for the optimal design also becomes more complicated. Fan pressure ratio (FPR), bypass ratio (BPR), overall pressure ratio (OPR) and combustor outlet temperature (T4) at the design point must be optimized as in the case of the conventional engine. Additionally, T4 at the

Copyright ©2011 by the American Institute of Aeronautics and Astronautics Inc. All rights reserved

2 Intercooler Recuperator

Variable guide vanes Figure 1 NEWAC IRA engine architecture cruise phase, the setting of the LPT nozzle can also be varied to find the optimal engine performance. Effects of varying these parameters have been previously discussed in [8][9]. In this study, these effects are further investigated and quantified. The design variables involved in this study are summarized in Table 1. Unless explicitly stated, all the parameters refer to the design point, which is the top of climb point (TOC). Notice that the influence of the intercooler auxiliary nozzle is reflected by the intercooler cold stream mass flow. Table 1 also presents the structure by which the design parameters have been grouped. A systematic parameter search is first carried out in FPR, fan mass flow and core mass flow to complete step I. The optimal point determined by this search and all other parameters set to nominal values are then carried over to Step II which optimizes intermediate pressure compressor (IPC) pressure ratio and high pressure compressor (HPC) pressure ratio. The same approach is then employed in Step III, Step IV and Step V. Design space limits and constraints The selecting of the parameters listed in Table 1 is limited to certain bounds due to several reasons. Firstly, the combined design parameters must be able to

Parameter Fan pressure ratio Fan inlet mass flow Core mass flow IPC pressure ratio HPC pressure ratio Cooling mass flow Combustor outlet temperature at cruise Intercooler cold stream mass flow Intercooler effectiveness LPT stator area at cruise Temperature change across recuperator cold surface

Abbr. FPR W1

Step I

W23 πIPC πHPC

II

W41 T4CR

III

W132 εIC

IV

A5CR ΔTREC

V

Table 1 Engine design variables and grouping of optimization parameters produce a feasible engine cycle. For example, with a fixed core size, the selected FPR and BPR must be within the range so that the power consumption of the fan will not surpass the power production of the LPT. The performance code of the TERA2020 will automatically interrupt the execution when such a limitation is reached. Secondly, the engine sizing must be feasible with the selected design parameters. For example, later we can see that the optimizer tries to design a recuperator with very high effectiveness. However, such a recuperator cannot be built due to the limited installation space in the engine hot nozzle. This effect is implemented in the weight modeling of the TERA and thus this limit can automatically be detected. Similar limitations also apply to the intercooler. Besides the above mentioned limits, the selections of the design parameters are also constrained by the state of art technology level. For example, turbine blade metal

3 temperature cannot exceed certain values due to the materials used. In this study, to enable a fair comparison between the nominal engine and the optimal engine, the blade metal temperature levels for the main different points of the nominal engine were used as upper limits during the study. Another example of the constraints is the LPT outlet temperature. To ensure that the yield stress of the Inco625 material used for the recuperator is not exceeded, outlet temperatures from the LPT are limited to 922 K here [10]. Furthermore, the design parameters are constrained when the engine is installed on an aircraft and to meet certain aircraft performance requirements. For example, the thrust of the engine must be sufficient to meet the regulations on take-off distance and the climb time to a certain altitude. All the constraints discussed here are summarized in Table 2. Parameter FAR take-off distance Climb time to 35000 ft Combustor outlet temperature at take-off Combustor outlet temperature at cruise HPT blade metal temperature at take-off HPT blade metal temperature at top of climb HPT blade metal temperature at cruise LPT outlet temperature

Abbr. -

Bound < 2.5 [km] < 22.5 [min]

T40TO

< 1950 [K]

T40CR

< 1850 [K]

T41mTO

< 1337 [K]*

T41mTOC

< 1295 [K]

T41mCR

< 1256 [K]

T5

< 922 [K]

Table 2 Optimization Constraints *temperature outside the thermal barrier coating

Propulsor Optimization There are three parameters of the propulsor to vary: FPR, the fan mass flow and the core mass flow. For a conventional engine with a fixed core mass flow, an increasing fan mass flow will increase the BPR and reduce the specific thrust. This will increase the propulsive efficiency and the engine fuel burn will benefit from this. However, increasing the fan mass flow also means that the fan size is increased, which will increase the fan weight, the LPT weight and the nacelle drag. Thus, for a fixed core, there is an optimal value for the fan mass flow. Similarly, for a fixed fan size, a decreasing core size will achieve a better propulsive efficiency. But due to the limitation on take-off distance and climb time to altitude, there is a lower bound on the fan size. When the fan and core sizes are fixed, i.e., the BPR is fixed; there is an optimal FPR value. When the FPR increases, both the thermal efficiency and the propulsive efficiency increase. However, when the FPR increases further, the propulsive efficiency drops and the overall efficiency achieves a maximal value. Also, the increased FPR will increase the LPT weight and hence the total engine weight. Much of these trends can still be seen from the parametric results of the intercooled recuperated engine propulsor. For example, one can see from Figure 2 that the increasing FPR is beneficial to the fuel burn until it reaches an optimal value. This is clearer when the fan size is fixed and the core size is floating, as shown in Figure 3. Besides these common trends between a conventional engine and an IRA engine, the trade-off between propulsive efficiency, thermal efficiency and engine weight is more complicated for an IRA engine. For example, increasing the FPR

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Figure 2 Fuel burn variation with FPR and fan mass flow

Figure 4 Engine weight variation with FPR and fan mass flow

will consume more power delivered by the LPT. As a consequence, the outlet temperature level at the LPT is reduced. This makes the design of the recuperator more difficult, since now its effectiveness must be increased in order to recuperate a sufficient amount of energy from this flow. As the recuperator weight increases progressively with the required design effectiveness, the engine weight also increases considerably and offsets the benefits of the increasing FPR, which can be seen in Figure 4. A similar trend applies to the fan mass flow due to the same reason: an increased fan size will consume more power and leave less for the recuperator. Here, the recuperator size is not constrained for the purpose of visualizing the trade-

off factor, thus the engine weight gets excessive in the upper right corner of the contour, i.e. when a fan with larger diameter and larger fan pressure ratio is designed. From the above discussions one would draw the conclusion that compared to conventional engines, the propulsor parameters of the IRA engine probably have a larger influence on the whole engine performance.

Figure 3 Fuel burn variation with FPR and core mass flow

With the other parameters in Table 1 frozen, -3.25% of ΔFPR, +2.0% of ΔW1 and -9.0% of ΔW23 compared to the nominal engine can bring 1.78% of fuel saving. The close to optimal point is represented by the square in Figure 2. As a matter of fact, the parametric study results show that the engine should have an even smaller specific thrust with a smaller core size. However, a large number of these designs were ruled out by the constraints of high pressure turbine (HPT) metal temperature at cruise, as shown by the dashed lines in Figure 2. Also, the regulation on climb time to certain altitude limits the design space, as shown by the solid lines in Figure 2. When one looks at these results again, it is clear that promising engines designs are found among those with a small core having a high BPR. For these engines the corresponding optimal FPR has to decrease slightly from

5

Figure 5 Fuel burn variation with IPC pressure ratio and HPC pressure ratio the nominal point, which can also be seen in Figure 3. Optimization of the IPC and HPC With the FPR, W1 and W23 set to the optimal value obtained from the last step, the IPC pressure ratio and the HPC pressure ratio can be optimized to achieve further fuel saving. The block fuel contour with respect to these variables is shown in Figure 5. The dashed lines in the figure show the HPT blade metal temperature at cruise. The results show that the IPC should have higher pressure ratio while the HPC should decrease the pressure ratio slightly. To understand the influence of these two parameters, the fuel burn is again broken down into thermal efficiency, propulsive efficiency and engine weight. The two design parameters involved here have little influence on the propulsive efficiency. However, both thermal efficiency and engine weight are strongly influenced when IPC pressure ratio and HPC pressure ratio change. Thus the trade-off is mainly between thermal efficiency and the engine weight. The thermal efficiency of an ideal recuperated cycle is depending on the overall pressure ratio (OPR) and the temperature ratio between the combustor outlet and the compressor

Figure 6 Thermal efficiency variation with OPR inlet. Increasing the IPC pressure ratio and HPC pressure ratio will increase the OPR, thus it will increase the thermal efficiency and reduce the fuel consumption. However, once the optimal value of the OPR is reached, the thermal efficiency will drop. This is also illustrated in Figure 6, which is extracted from the same parametric study as shown in Figure 5. It is to be noticed that the optimal value is much lower than the conventional engines, which is in line with the previous results [8]. Thus, when the IPC pressure ratio increases to some extent, it is necessary to decrease the HPC pressure ratio to maintain the optimal OPR. The influence of the OPR on engine weight is roughly the opposite of its influence on the thermal efficiency. An increasing OPR will increase the compressor outlet temperature and the cold side of the recuperator gets hotter. Meanwhile, more energy is extracted from the core flow and the LPT outlet temperature decreases. The hot side of the recuperator gets cooler. Since the recuperator is designed to heat the air with a fixed amount of temperature increase, much more heat exchange surface is needed due to this reduced temperature gradient and

6 this makes the engine unacceptably heavy. The split of the IPC and HPC pressure ratio at design point does not have strong influence on the fuel burn. However, the product of the FPR and IPC pressure ratio is kept to be roughly equal to the HPC pressure ratio. Shifting the pressure ratio split more towards the IPC will increase both the exit temperature from the IPC and the HPC. This will decrease the temperature gradient across the recuperator and hence decrease the engine thermal efficiency. On the other hand, if more pressure ratio is put on the HPC, the effect of intercooling on the cycle becomes smaller and the HPT has to deliver more power to the HPC. As a consequence, the exit temperature from the LPT drops as well and the temperature gradient across the recuperator is reduced. For these reasons, there is an optimal point to introduce the intercooling into the cycle. With the values from step I frozen, optimization results show that -4.75% of ΔπHPC and +9.25% of ΔπIPC can achieve 2.36% fuel saving compared to the nominal point, which corresponds to an additional 0.58% compared to step I. The optimal point is represented by the square in Figure 5. The results also indicate that the IPC pressure ratio should increase further with a decreasing HPC pressure ratio. However, such design combinations will fail to meet the constraints on the climb time constraint. Engine Rating at Cruise and Cooling For conventional engines, the combustor outlet temperature T4 at the cruise phase is determined by its design point value and the cruise phase thrust requirement. Due to the introduction of the variable nozzle area in the LPT, however, T4 at the cruise phase can

Figure 7 Fuel burn variation with HPT cooling mass flow and T4 at cruise be selected freely in the IRA engine. Due to the utilization of the recuperator, the increasing T4 at cruise can improve the thermal efficiency of the engine and achieve a better fuel consumption. At the same time, the propulsive efficiency will decrease with the higher T4 at cruise. However, before this optimum value could be reached, there are upper bounds for this temperature due to the consideration of combustor design and turbine blade life as well as NOx level. In this study T4 at cruise is limited to 1850 K. It is worth noticing that since there the engine sizing is carried out at the take-off and top of climb, change of the temperature at cruise phase has no impact on the engine weight, nor the nacelle drag. Thus, the trade-off is mostly between the thermal efficiency and the propulsive efficiency. The fuel burn contour is shown in Figure 7. With the values from step I and II fixed, +1.5% of ΔT4CR and +8.85% of ΔW41 will give a fuel burn saving of 3.00% compared to the nominal engine, which corresponds to an additional 0.64% compared to step II. Such a design combination meets the constraints on HPT metal temperature and climb time regulation, as shown by the solid

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Figure 8 Fuel burn variation with intercooler cold stream mass flow and intercooler effectiveness and dashed lines in Figure 7. Intercooler Effectiveness and Cold Stream Mass Flow Rate With increasing intercooler effectiveness the thermal efficiency of the engine will decrease and the fuel burn will go up, as can be seen from Figure 8. Also, a reduced cold stream mass flow can improve the thermal efficiency which is an off-design effect from the engine performance. However, when this parameter is decreased too much the compressor outlet temperature will become too high and there will be no temperature gradient left for the recuperator to achieve the desired temperature drop. The thermal efficiency contour has the same trend as shown in Figure 8, but the engine weight has the opposite trend (not shown). This is an indication that the thermal efficiency dominates the fuel burn result. The engine weight is not as sensitive as in the other cases studied in the previous steps. The reason behind this is that the intercooler weight and the recuperator weight compromise each other. A higher intercooler effectiveness will result in a heavier intercooler installation. However, it also creates a lower HPC outlet temperature and thus a

Figure 9 Fuel burn variation with LPT nozzle area at cruise and ΔT across recuperator cold side higher temperature gradient for the recuperator. This will effectively reduce the recuperator size and thus the total engine weight sensitivity. The design with minimal fuel burn shows a reduction of a -56.4% ΔW132 with a -0.8% ΔεIC at top of climb. This will decrease the fuel burn by 1.42% compared to the results from the last step. Thus it achieves a fuel burn saving of 4.42% compared to the nominal engine. LPT nozzle area / Recuperator Effectivness The increasing recuperator temperature is beneficial to the fuel burn, as can be seen from Figure 9. An increase of the temperature change across the efficiency increases. However, the increasing ΔTREC will also cause the recuperator weight to increase, thus there will be an optimal value for this design parameter. The variation of the LPT nozzle area will affect engine performance and more importantly, the thermal efficiency. Again this is because the ΔTREC at cruse will vary. With the values from all the previous steps fixed, +4.5% of ΔA5CR and +2.75% of ΔTREC will give a fuel burn of -4.61% compared to the nominal engine. This corresponds to an

8 additional 0.19% of fuel saving compared to the previous step. From Figure 9 it is to notice that large design space is ruled out the constraint HPT metal temperature at cruise, as shown by the dashed line. Summary and conclusions The design parameters and its influence on fuel burn are summarized in Table 3. From the results it can be seen that an optimal IRA design should have a smaller core size and a reduced specific thrust. The results also show that the Abbreviation Optimal value ΔFPR -3.25% Step I ΔW1 2.00% ΔW23 -9.00% +9.25 ΔπIPC Step II -4.75 ΔπHPC ΔT4CR +1.50% Step III ΔW41 +8.85% ΔW132 -56.4% Step IV ΔεIC -0.80% ΔTREC +2.75% Step V ΔA5CR +4.50% Fuel burn -4.61% compare to nominal Table 3 Summary of changes in design parameters with step wise parametric study recuperator should be designed to achieve higher effectiveness compared to the nominal design. Such a design requires the recuperator to have a bulky installation and makes the already difficult integration task of the recuperator even more challenging. The described work takes a first step towards a full optimization of the IRA power plant. Influence and interaction of different design parameters on propulsive efficiency, thermal efficiency, engine weight and nacelle drag is analyzed. Some better understanding of the IRA engine design space is

gained through such an analysis. Experience also shows that result from such a parametric study is close to the result of the full optimization. Acknowledgment This study has been performed in the NEWAC project under the European Commission Contract No. AIP-CT-2006-030876. The authors gratefully acknowledge this funding. References [1] McDonald, C., Massardo, A., Rodgers, C. and Stone, A., “Recuperated gas turbine aeroengines, part I: early development activities”, Aircraft Engineering and Aerospace Technology: An International Journal, Vol 80, No.2, pp139-157, 2008 [2] McDonald, C., Massardo, A., Rodgers, C. and Stone, A., “Recuperated gas turbine aeroengines, part II: engine design studies following early development testing”, Aircraft Engineering and Aerospace Technology: An International Journal, Vol 80, No.3, pp280-294, 2008 [3] McDonald, C., Massardo, A., Rodgers, C. and Stone, A., “Recuperated gas turbine aeroengines, part III: engine concepts for reduced emissions, low fuel consumption, and noise abatement”, Aircraft Engineering and Aerospace Technology: An International Journal, Vol 80, No.4, pp408-426, 2008 [4] Wilfert, G., Sieber, J., Rolt, A., Baker, N., Touyeras, A. and Colantuoni, S., “New Environmental Friendly Aero Engine Core Concepts”, ISABE-2007-1120

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