Design point performance results

Figure 3.1 shows the results obtained for specific thrust and specific fuel consumption, respectively, as a function of compressor pressure ratio for several values of bypass ratio. This design point analysis was carried out for an altitude ofh= 10000mand flight Mach number ofM₀= 0.8. Concerning the level of technology, this analysis is related with level 4; the design parameters simulated were a fan pressure ratio ofπf = 1.6 and a turbine inlet temperature of Tt4 = 1600K. The coupling between turbine and compressor was assumed to have a efficiency of ηm = 0.99. For the specific case of a bypass ratio B = 5, the same results for specific thrust and specific fuel consumption, respectively, are compared with those produced by the packageGasTurb^R. Good agreement is obtained between the two matching sets of results, specially for low values of compressor pressure ratio. The mean relative errors between the results presented here and those produced by the packageGasTurb^R for specific thrust and specific fuel consumption are 3.8% and 3.1%, respectively.

(a)Ψas a function ofπc (b)Sas a function ofπc

(c)Ψas a function ofπc:GasTurb^R comparison (d)Sas a function ofπc:GasTurb^R comparison

Figure 3.1: Design point results as a function of compressor pressure ratioπc

Figure 3.2 shows the results obtained for specific thrust and specific fuel consumption, respectively, as a function of flight Mach number for several values of bypass ratio. This design point analysis was carried out for an altitude ofh= 10000m and compressor pressure ratio ofπc = 30. Concerning the level of technology, this analysis is related with level 4 and the design parameters simulated were a fan pressure ratio ofπ_f = 1.6and a turbine inlet temperature ofT_t4= 1600K. The coupling between turbine and compressor was assumed to have a efficiency ofη_m= 0.99. For the specific case of a bypass ratio B = 5, the same results for specific thrust and specific fuel consumption, respectively, are compared with those produced by the packageGasTurb^R. Very good agreement is obtained between the two matching sets of results, except for the supersonic range, which is out of the operating conditions of interest for the present study. The mean relative errors between the results presented here and those produced by the packageGasTurb^R for specific thrust and specific fuel consumption at subsonic conditions are 1.8%

and 1.2%, respectively.

(a)Ψas a function ofM0 (b)Sas a function ofM0

(c)Ψas a function ofM0:GasTurb^R comparison (d)Sas a function ofM0:GasTurb^R comparison

Figure 3.2: Design point results as a function of flight Mach numberM0

Figure 3.3 shows the results obtained for specific thrust and specific fuel consumption, respectively, as a function of fan pressure ratio for several values of bypass ratio. This design point analysis was carried out for an altitude ofh = 10000mand flight Mach number ofM0 = 0.8. Concerning the level of technology, this analysis is related with level 4; the design parameters simulated were a compressor pressure ratio ofπ_c = 30and a turbine inlet temperature ofT_t4= 1600K. The coupling between turbine and compressor was assumed to have a efficiency of η_m = 0.99. For the specific case of a bypass ratioB = 5, the same results for specific thrust and specific fuel consumption, respectively, are com- pared with those produced by the packageGasTurb^R. Good agreement was obtained between the two matching sets of results, with results following the same trend. The mean relative errors between the results presented here and those produced by the package GasTurb^R for specific thrust and specific fuel consumption are 4.2% and 6.3%, respectively.

(a)Ψas a function ofπf (b)Sas a function ofπf

(c)Ψas a function ofπf:GasTurb^R comparison (d)Sas a function ofπf:GasTurb^R comparison

Figure 3.3: Design point results as a function of fan pressure ratioπ_f

Figure 3.4 shows the results obtained for specific thrust and specific fuel consumption, respectively, as a function of bypass ratio for several values of fan pressure ratio. This design point analysis was carried out for an altitude ofh = 10000mand flight Mach number ofM0 = 0.8. Concerning the level of technology, this analysis is related with level 4; the design parameters simulated were a compressor pressure ratio ofπ_c = 30and a turbine inlet temperature ofT_t4= 1600K. The coupling between turbine and compressor was assumed to have a efficiency ofη_m= 0.99. For the specific case of a fan pressure ratioπ_f = 1.5, the same results for specific thrust and specific fuel consumption, respectively, are com- pared with those produced by the packageGasTurb^R. Again, good agreement is obtained between the two matching sets of results for all the evaluated spectrum of bypass ratios. The mean relative errors between the results presented here and those produced by the packageGasTurb^R for specific thrust and specific fuel consumption are 4.6% and 6.1%, respectively.

(a)Ψas a function ofB (b)Sas a function ofB

(c)Ψas a function ofB:GasTurb^R comparison (d)Sas a function ofB:GasTurb^R comparison

Figure 3.4: Design point results as a function of bypass ratioB

Figure 3.5 shows the result obtained for fuel/air ratio, as a function of compressor pressure ratio in comparison with those produced by the packageGasTurb^R. This design point analysis was carried out for an altitude ofh= 10000mand flight Mach number ofM0= 0.8. Concerning the level of technology, this analysis is related with level 4; the design parameters simulated were a fan pressure ratio ofπf = 1.6 and a turbine inlet temperature ofT_t4 = 1600K. The coupling between turbine and compressor was assumed to have a efficiency ofη_m= 0.99. For the fuel/air ratio, it is possible to see that yet again there is a good agreement, despite the fact that the scaling of the figure enhances the difference between the two results. The mean relative error between the results presented here and those produced by the packageGasTurb^R is 6.7%.

Figure 3.5: Fuel/air ratiof as a function ofπc

From Figure 3.1 and Figure 3.5 it is possible to understand the influence of compressor pressure ratio and bypass ratio on engine performance. With increased values of bypass ratio, specific thrust and specific fuel consumption decrease; however, the rate of decrease of these performance parameters is not proportional to the rate of increase of bypass ratio values. Concerning the compressor ratio influence, it can be seen that with its increase, specific thrust increases untilπ_c ≈8 and then it stays approximately stable; specific fuel consumption decreases more sharply for lower values of compressor pressure ratio, and more moderately for higher values. Fuel/air ratio decreases with the increase of compressor pressure ratio.

Figure 3.2 depicts the influence of flight Mach number and bypass ratio. It shows a strong depen- dency between specific thrust and flight Mach number: as the latter increases, the value of specific thrust decreases; the same happens with bypass ratio (the higher the bypass ratio is, the lower the specific thrust is). For specific fuel consumption, this influence is not so strong, except in supersonic conditions.

However, it is obviously possible to see that the greater the flight Mach number is, the higher the fuel consumption will be.

The effect of fan pressure ratio and bypass ratio is shown in Figure 3.3. There we can observe that, to all bypass ratios, there is a maximum in engine performance (highest specific thrust and lowest specific fuel consumption): such point indicates an optimum fan pressure ratio. For maximizing specific thrust, bypass ratio should be at the lowest; on the other hand, for minimizing specific fuel consumption,

the highest bypass ratio is the best.

Another design parameter studied in this section is the influence of bypass ratio. In Figure 3.4 it is possible to evaluate how a variation in performance correlates with a variation in bypass ratio and fan pressure ratio. Again, it is possible to observe that, to all fan pressure ratios, there is an optimum bypass ratio (easily seen in the specific fuel consumption chart).

Figure 3.6 shows the results obtained for specific thrust, as a function of specific fuel consumption for several values of turbine inlet temperature and compressor pressure ratio, using different levels of technology. This design point analysis was carried out for an altitude ofh= 10000mand flight Mach number ofM0 = 0.8. Concerning design parameters, a fan pressure ratio ofπf = 1.6and a a bypass ratio ofB= 5were introduced; the coupling between the turbine and compressor was assumed to have an efficiency ofηm= 0.99.

(a)1^stlevel of technology (b)2^ndlevel of technology

(c)3^rdlevel of technology (d)4^thlevel of technology

Figure 3.6: Design point results as a function of turbine inlet temperatureTt4and compressor pressure ratioπc

Combining the effects of compressor pressure ratio and inlet turbine temperature it is possible to plot charts as in Figure 3.6. Note that specific thrust and fuel consumption increase with increasing turbine entry temperature and decrease with increasing compressor pressure ratio. By analysing the figure, it is possible to find a compressor pressure ratio that combines the main goals of design point

analysis: setting the specific thrust at its maximum and simultaneously keeping fuel consumption as low as possible. It is also possible to understand the effect different levels of technology have in specific thrust, although it shows mainly in specific fuel consumption.

There is a good agreement between the results obtained usingGasTurb^R and those obtained here.

The existing differences can be traced back to the fact thatGasTurb^R software implements a model that is more robust than the one used here. Thus, it contains less simplifications. On the other hand, several components of the engine are there characterized differently.