. 2 . 4 . 6 . 8 1.0 1.2 1.4 1.6 1.8 2 . 0 2 . 2 2 . 4 MACH NUMBER
Figure 7.18 Augmentation in Order to Satisfy Energy and Maneuverability Requirement (P$)
I-22 PRELIMINARY ENGINE DESIGN
After defining a configuration based on previous experi-ence, the anticipated technology level will define critical cycle parameters. The cycle variables directly impacted by technology level are: turbine inlet temperature (T41), compressor discharge temperature (T3), component effi-ciency levels, turbine cooling flows, and exhaust nozzle cooling flows. For a mixed flow turbofan, selecting the exhaust stream mixer, augmentor, and exhaust nozzle type completes the technology definition. Component matching criteria, inlet airflow schedules, minimum al-lowable fan and compressor stall margins, overspeed ca-pability, etc. are all specified to complete the cycle definition. Having established these limits and cycle guidelines, the primary remaining unknowns are the de-sign bypass ratio, and the fan and compressor pressure ratio. To facilitate this selection, a matrix of design by-pass ratio/fan pressure ratio is constructed which is con-sistent with the cycle limits and matching guidelines.
Figure 1.19 represents such a matrix for a family of en-gines with overall cycle pressure ratios between 18 and 30. Fan pressure ratios range from 3.6 to 4.3, and core compressor pressure ratios range from 5 to 7. These ranges were established by adding or subtracting a stage
to a reference compressor and by the preference for a single stage low pressure turbine. This is a consideration but not a constraint as the selection of the design bypass ratio will also impact the required turbine stage count.
The following observations can be made from the figure.
For a given core pressure ratio, increasing the design fan pressure ratio will increase the overall cycle pressure ra-tio. The higher overall pressure ratio engines are re-stricted to lower flight Mach numbers as a result of the T3 limit. However, the T3 limit is reached at a lower ram temperature which allows the core size to be re-duced (bypass ratio increased). Conversely, the lower overall pressure ratio engines can attain higher flight Mach numbers before the T3 limit is reached; but the higher ram temperature requires more core energy to drive the fan, and the core size must be increased for the specified T41 limit (bypass ratio reduced). If the overall pressure ratio is held constant (constant T3 and design flight Mach number), the fan pressure ratio has a large impact on the design bypass ratio. However, if the over-all pressure ratio is over-allowed to vary, fan pressure ratio has a much smaller impact on design bypass ratio.
0.9
DESIGN OVERALL PRESSURE RATIO
30 24 ! 8 6 1.8 2 . 0 2.2 2.4 2 . 6
DESIGN F L I G H T MACH NUMBER 2 . 8
Figure 1.19 Cycle Matrix
PRELIMINARY ENGINE DESIGN 1-23
For the cycles defined by the matrix, thrust ratios and specific fuel consumption are used to determine the rela-tive performance between the cycles and to provide a ba-sis for making a selection. Thrust ratios are defined as the ratio of thrust at a given flight condition to sea level static maximum augmented thrust.
Figure 1.20 presents thrust ratios for maximum dry op-eration at supersonic Mach numbers. The lower overall pressure ratio cycles have the highest thrust ratios be-cause of their ability to maintain airflow out to higher Mach numbers. If, for example, a dry thrust ratio of 0.4 were required at Mach 2.15, a fan pressure ratio of four and a core pressure ratio of six would represent the most appropriate cycle for satisfying this requirement.
Figure 1.21 shows the maximum augmented thrust ra-tios for the matrix at the same supersonic Mach num-bers. Again, the lower overall pressure ratio engines have greater thrust ratios. Furthermore, as fan pressure ratio is reduced (design bypass ratio increased), the en-gines have an increasing augmented thrust capability to progressively higher Mach numbers. The impact of the augmentor is reduced for lower bypass ratio engines
be-cause a larger percentage of the air in the augmentor has already been heated in the main combustor, reducing the oxygen content and temperature rise capability of the augmentor.
The effect of design overall pressure ratio and design by-pass ratio on minimum SFC is shown in Figure 1.22.
The higher bypass and higher overall pressure ratio cy-cles have the best SFC. Therefore, selecting the cycle with the highest overall pressure ratio consistent with the maximum flight Mach number and the maximum bypass ratio that satisfies the critical thrust ratio would provide the best uninstalled SFC. However, it should be noted that providing the minimum SFC at a rated thrust does not necessarily result in maximizing aircraft range. If an engine were designed to have minimum SFC at a sub-sonic cruise point, any changes in aircraft gross weight or drag would shift the resulting system away from the minimum SFC point. In addition, aircraft gross weight and L/D changes throughout the cruise segment of the flight as on-board fuel is consumed. Consequently, it is important that an engine have a relatively flat SFC char-acteristic at key cruise conditions (Figure 1.23).
THRUST
REQUIRED THRUST RATIO
1.6 2.0 2.4
FLIGHT MACH NUMBER
2.6
Figure 1.20 Maximum Dry (Non-Augmented) Supersonic Thrust Ratios
1-24 PRELIMINARY ENGINE DESIGN
THRUST THRUST
SLS
1.2
1.1'
1.0
0.9
0.8
0.7
0.6
0.0
365
CONSTANT OVERALL PRESSURE RATIO
1.6 1.8 2.0 2.2 2.4 2.6 2.8 FLIGHT MACH NUMBER
Figure 1.21 Maximum Augmented Supersonic Thrust Ratios
x A
SFC
4,
0.
•2
- 4 ,
-6
CPR - 5
FPR - 3.6
CONSTANT OVERALL PRESSURE RATIO
18 20 22 24 26 28 DESIGN OVERALL PRESSURE RATIO
Figure 1.22 Minimum Specific Fuel Consumption for Subsonic Cruise 30
PRELIMINARY ENGINE DESIGN 1-25
-—DRY PUR—
Figure 1.23 Maximum subsonic range does not necessarily occur at SFC minimum. An engine which has been sized by high Mach thrust or maneuverability requirements will be throttled well back from SFC minimum at cruise. An aircraft with a long subsonic cruise leg will be designed to cruise at wing UD (lift to