Ultra-High Bypass Ratio Turbofan for Next-Generation Large Aircraft

Nicholas Turo-Shields, Sebastian Perkinson, Shashank Kashyap, Chris Vodney, Maunik Patel, Gerardo Martinez, Aditya Anilkumar

AAE/ME 538 14 December 2018 Table of Contents

Nomenclature 3

1 Abstract 4

2 Design Requirements 4

3 Baseline Engine 4

4 Team Structure 8

5 Design Methodology 8 5.1 Choice of Architecture ...... 8 5.2 BPR and Core Sizing ...... 9 5.2.1 Initial Core Scaling (constant fan mass flow) ...... 9 5.2.2 Bypass Scaling (constant core scale) ...... 9 5.2.3 Evaluation of Design for Takeoff ...... 10 5.2.4 Sizing for Cruise ...... 11 5.2.5 Further Improvements ...... 12 5.3 Gearbox ...... 12 5.4 Compressors ...... 13 5.4.1 LP Compressor (Fan) ...... 13 5.4.2 IP Compressor (Booster) ...... 15 5.5 LP Turbine ...... 16 5.5.1 Optimization of Specific Fuel Consumption ...... 17 5.5.2 Optimization of Core Flow ...... 17 5.5.3 Final Results for LP Turbine ...... 18 5.6 Materials ...... 19 5.7 Engine Cycle Improvements ...... 21 5.8 Risks and Concerns ...... 22

6 Our Engine Solution 23

7 Off-Design Analysis 24 7.1 Off-Design Operating Point ...... 24 7.2 Mission Profile ...... 25 7.3 Flight Envelope ...... 26 7.4 Mission Points ...... 28 7.5 Results ...... 28

8 Trade-oﬀ Study 30 8.1 Parametric Study ...... 31 8.2 Baseline Engine ...... 31 8.3 Geared Solution ...... 32

9 Limits of Analysis 33

10 Appendix A - Velocity Triangles 34

11 Appendix B - GasTurb Engine Conﬁguration (Final GTF) 35

12 Appendix C - Material Properties 40

2 Symbols and Abbreviations

A Area (m2) L Low-pressure spool alt Altitude (m) LP- Low-pressure amb Ambient LPC LP compressor (fan) ax Axial LPT LP Turbine Bld Bleed M Mach number BPR Bypass ratio N Spool speed corr Corrected NGV Nozzle guide vane (of a turbine) C Compressor o Outer Cl Cooling P Total pressure (kPa) d Diameter (m) prop Propulsion dH Enthalpy difference R Gas constant dp Design point rel Relative f Fuel RNI Reynolds number index far Fuel-air-ratio s Static FN Net thrust (kN) S NOx NOx severity parameter h Enthalpy SFC g H High-pressure spool Specific fuel consumption kN·s HdlBld Handling bleed t Tip (blade) or time HP- High-pressure (compressor, turbine) T Total temperature (K) i Inner U Blade (tip) velocity (m/s) IP- Intermediate-pressure V Velocity (m/s) IPC IP compressor (booster) W Mass flow (also denotedm ˙ ) (kg/s)

Station Designations

0 Ambient 41 First turbine stator exit = rotor inlet 1 Engine inlet 42 HPT exit before addition of cooling air 2 First compressor inlet 43 HPT exit after addition of cooling air 21 Inner stream fan exit 44 IPT inlet 13 Outer stream fan exit 45 IPT stator exit 16 Bypass exit 46 IPT exit before addition of cooling air 18 Bypass nozzle throat 47 IPT exit after addition of cooling air 24 IP compressor exit 48 LPT inlet 25 HP compressor inlet 49 LPT exit before addition of cooling air 3 HPC exit, cold side heat exchanger inlet 5 LPT exit after addition of cooling air 31 Burner inlet 8 Nozzle throat 4 Burner exit

3 1 Abstract

Boeing and Airbus are considering replacement engines for their 787, A380, and A350 airplanes. Newer engine technologies enable the core to operate more efficiency and deliver more power. Higher bypass ratio engines are being considered to improve propulsive efficiency. An engine model has been provided. The task is to keep the same core but design an improved LP/IP stage that utilizes the existing core to improve the propulsive efficiency over the mission of the aircraft. The performance characteristics and total fuel consumption should be estimated over the mission. Attention should be paid to weight, dimensions, stage aerodynamic compatibility, and technical feasibility (Materials, etc). Operating cost and maintenance cost (limited stage count, reduced blade count) should be considered.

2 Design Requirements

• Select an engine architecture (1/2/3-spool, geared, etc.)

• Flight Design Points – Takeoﬀ: sea-level – Cruise: 12,190 m at Mach 0.85 – Range: 17,600 km

• Takeoﬀ Thrust ≥ 374 kN

• Takeoﬀ Power ≥ 200 kW

• Overall Pressure Ratio (OPR, p3/p2): 60

• Turbine Inlet Temperature (T4): 1930 K

• Power oﬀ-take: 200 kW

• If a geared design is chosen, assume a mass of 0.00482768 kg/kW

3 Baseline Engine

The baseline engine we are comparing our redesigned engine against is the Rolls-Royce Trent XWB.

Table 1: Basic design features of the baseline engine.

Engine Type Axial, turbofan Number of fan/booster/compressor stages 1, 8, 6 Number of HP/IP/LP turbine stages 1, 2, 7 Combustor type Annular Maximum net thrust at sea-level 396 kN g SFC at cruise at Mach 0.85 & 12.19 km altitude 18.2 kN·s Overall pressure ratio at max. power 50 Bypass ratio 9.3 Max. envelope diameter 2.997 m Max. envelope length 4.064 m Dry weight less tail-pipe 5,445 kg Turbine Inlet Temperature 1784 K

4 Table 2: Input to GasTurb of the baseline engine model.

Property Unit Value Comment Intake Pressure Ratio 1 No (0) or Average (1) Core dP/P 1 Inner Fan Pressure Ratio 1.4 Outer Fan Pressure Ratio 1.43 Core Inlet Duct Pressure Ratio 1 IP Compressor Pressure Ratio 6.3 Compressor Interduct Pressure 0.985 Ratio HP Compressor Pressure Ratio 5.76 Bypass Duct Pressure Ratio 0.975 Inlet Corr. Flow W2Rstd kg/s 1442.92 Inlet corrected flow rate standard day Design Bypass Ratio 9.3 Burner Exit Temperature K 1783 Burner Design Efficiency 0.9995 Burner Partload Constant 1.6 Used for off design only Fuel Heating Value MJ/kg 43.124 Overboard Bleed kg/s 0 Power Offtake kW 50 HP Spool Mechanical Efficiency 0.99 IP Spool Mechanical Efficiency 0.999 LP Spool Mechanical Efficiency 0.999 Burner Pressure Ratio 0.96 Ipt Interd. Ref. Press. Ratio 0.992 Lpt Interd. Ref. Pressure Ratio 1 Turbine Exit Duct Press Ratio 0.99

Figure 1: Baseline engine station diagram and ﬂows, 3-spool.

5 Figure 2: Baseline engine model schematic, 3-spool.

Table 3: GasTurb summary of the baseline model engine at sea-level.

W T P WRstd Station kg/s K kPa kg/s FN = 396.03 kN amb 288.15 101.325 TSFC = 7.8416 g/(kN*s) 2 1442.919 288.15 101.325 1442.920 WF = 3.10549 kg/s 13 1302.830 322.15 144.895 963.323 s NOx = 2.28101 21 140.089 320.77 141.855 105.575 BPR = 9.3000 22 140.089 320.77 141.855 105.575 Core Eff = 0.5278 24 140.089 562.99 893.687 22.201 Prop Eff = 0.0000 25 140.089 562.99 880.281 22.539 P3/P2 = 50.041 3 131.684 921.84 5070.420 4.707 P2/P1 = 1.00000 31 116.274 921.84 5070.420 P22/P21 = 1.00000 4 119.380 1783.00 4867.603 6.182 P25/P24 = 0.98500 41 126.384 1739.15 4867.603 6.463 P4/P3 = 0.96000 42 126.384 1406.66 1706.924 P44/P43 = 0.99200 43 134.789 1378.42 1706.924 P48/P47 = 1.00000 44 134.789 1378.42 1693.269 P6/P5 = 0.99000 45 139.693 1359.02 1693.269 18.154 P16/P13 = 0.97500 46 139.693 1156.75 785.978 P16/P6 = 0.69135 47 141.794 1151.55 785.978 P5/P2 = 2.03708 48 141.794 1151.55 785.978 36.542 V18/V8,id= 0.43099 49 141.794 857.50 206.407 A8 = 0.51485 m² 5 143.195 855.67 206.407 121.132 A18 = 4.83610 m² 8 143.195 855.67 204.343 122.356 XM8 = 1.00000 18 1302.830 322.15 141.272 988.024 XM18 = 0.70583 Bleed 0.000 921.84 5070.420 WBld/W2 = 0.00000 ------Efficiencies: isentr polytr RNI P/P CD8 = 1.00000 Outer LPC 0.9103 0.9147 1.000 1.430 CD 18 = 0.92331 Inner LPC 0.8900 0.8951 1.000 1.400 PWX = 50.00 kW IP Compressor 0.8991 0.9210 1.233 6.300 WlkLP/W25= 0.00000 HP Compressor 0.9290 0.9430 3.912 5.760 WBld/W25 = 0.00000 Burner 0.9995 0.960 Loading = 100.00 % HP Turbine 0.9094 0.8989 5.888 2.852 e442 th = 0.87880 IP Turbine 0.9061 0.8980 2.722 2.154 WCHN/W25 = 0.05000 LP Turbine 0.9193 0.9059 1.528 3.808 WCHR/W25 = 0.06000 ------WCIN/W25 = 0.03500 HP Spool mech Eff 0.9900 Nom Spd 1199 9 rpm WCIR/W25 = 0.01500 IP Spool mech Eff 0.9990 Nom Spd 5113 rpm WCLR/W25 = 0.01000 LP Spool mech Eff 0.9990 Nom Spd 2472 rpm ------hum [%] war0 FHV Fuel 0.0 0.00000 43.124 Generic

6 Table 4: GasTurb summary of the baseline model engine at cruise conditions.

W T P WRstd Station kg/s K kPa kg/s FN = 57.90 kN amb 216.65 18.760 TSFC = 18.1933 g/(kN*s) 2 492.903 248.02 30.097 1539.549 WF = 1.05338 kg/s 13 447.947 279.78 41.233 1084.669 s NOx = 1.18576 21 44.956 278.46 40.456 110.684 BPR = 9.9642 22 44.956 278.46 40.456 110.684 Core Eff = 0.5791 24 44.956 520.79 265.829 23.037 Prop Eff = 0.8113 25 44.956 520.79 261.536 23.415 P5/P2 = 2.21582 EPR 3 42.258 882.93 1629.562 4.600 P2/ P1 = 1.00000 31 37.313 882.93 1629.562 P22/P21 = 1.00000 4 38.367 1792.32 1567.317 6.186 P25/P24 = 0.98385 41 40.614 1746.30 1567.317 6.464 P4/P3 = 0.96180 42 40.614 1414.47 551.684 P44/P43 = 0.99206 43 43.312 1383.78 551.684 P48/P47 = 1.00000 44 43.312 1383.78 547.305 P6/P5 = 0.99001 45 44.885 1363.03 547.305 18.073 P16/P13 = 0.96831 46 44.885 1162.65 255.689 P16/P6 = 0.60473 47 45.559 1156.85 255.689 P5/P2 = 2.21582 48 45.559 1156.85 255.689 36.175 V18/V8,id= 0.45171 49 45.559 866.77 66.689 A8 = 0.51485 m² 5 46.009 864.50 66.689 121.080 A18 = 4.83610 m² 8 46.009 864.50 66.023 122.302 XM8 = 1.00000 18 447.947 279.78 39.926 1120.173 XM18 = 1.00000 Bleed 0.000 882.93 1629.562 WBld/W2 = 0.00000 ------Efficiencies: isentr polytr RNI P/P CD8 = 1.00000 Outer LPC 0.7361 0.7476 0.355 1.370 CD18 = 0.96000 Inner LPC 0.7197 0.7312 0.355 1.344 PWX = 50.00 kW IP Compressor 0.8089 0.8512 0.416 6.571 WlkLP/W25= 0.00000 HP Compressor 0.9100 0.9285 1.275 6.231 WBld/W25 = 0.00000 Burner 0.9974 0.962 Loading = 281.88 % HP Turbine 0.9092 0.8988 1.886 2.841 e442 th = 0.88048 IP Turbine 0.9042 0.8961 0.877 2.141 WCHN/W25 = 0.05000 LP Turbine 0.9015 0.8855 0.495 3.834 WCHR/W25 = 0.06000 ------WCIN/W25 = 0.03500 HP Spool mech Eff 0.9900 Speed 11999 rpm WCIR/W25 = 0.01500 IP Spool mech Eff 0.9990 Speed 5219 rp m WCLR/W25 = 0.01000 LP Spool mech Eff 0.9990 Speed 2673 rpm ------hum [%] war0 FHV Fuel 0.0 0.00000 43.124 Generic

Figure 3: Baseline engine thermodynamic cycle.

7 4 Team Structure

Nicholas Turo-Shields IPC design & optimization, report compilation Chris Vodney BPR & core sizing Aditya Anilkumar Turbine optimization Gerardo Martinez Oﬀ-design & mission analysis Shashank Kashyap Oﬀ-design & mission analysis Maunik Patel LPC design & optimization Sebastian Perkinson Materials & cooling

5 Design Methodology

Since the aircraft spends the majority of its time and therefore burns the most fuel during the cruise phase, that was taken as the design point. Using the Mission tab, engine characteristics can be calculated at both takeoﬀ and cruise. With this, constraints spanning takeoﬀ and cruise at the same time can be considered, simplifying the design process.

The Rolls Royce Trent XWB engine was modelled in GasTurb 13 which is a three spooled engine, a similar performance Geared Turbofan engine with 2 spools was also modelled in GasTurb 13. Both types of engines are better suited for the high bypass ratio engines with less speciﬁc fuel consumption and fuel burn, which is primary target for any commercial aircraft.

5.1 Choice of Architecture 3-Spooled Turbofan Design A third spool allows fan to be driven by low pressure turbine and introduces intermediate stage for compressor and turbine. This allows both the fan and intermediate stage to operate at close to their optimal speeds, however this leads to increased weight. But this allows us to reach similar pressure ratio as 2 spool but with less stages.

Geared 2-Spool Turbofan Design A gearbox is added between the fan and LPT shaft so that the intake fan is spinning at a more favourable rate, as the rotational speed of the LPT shaft is much higher than the fan prefers. Although adding gearbox adds weight, it allows for longer fan blades (since the centripetal loading depends on rotational speed) that can lead to improvement in fuel eﬃciency, reduction in noise. There is considerable reduction in NOx emissions and reduction in a spool saves production costs of lot of components.

Figure 4: Reduction gearbox used on the PW1100G [8].

The geared turbofan engine is not a very common choice of architecture and to justify this choice of selection, a 3 spooled engine was designed alongside geared turbofan. It was found geared turbofan has better performance

8 for the higher bypass ratio. Looking at the greater advantage of geared turbofan a new engine cycle was created. This geared turbofan was then followed by improvements to each component of the turbofan, as the constraint was not to modify core eﬃciency other key components like the Low pressure Turbines and Compressors along with Nacelle and Material improvements were considered to get a better performance.

The ﬁnal results of each component was combined as one engine and a ﬁnal engine design was proposed.

5.2 BPR and Core Sizing The design process started with a modified baseline engine to meet overall pressure ratio, power off-take, and turbine inlet temperature requirements. Both the 3-spooled and geared architectures were walked through the same parametric studies until justification was found that the geared engine would be required to achieve the best performance. The following sections discuss the key parametric studies used for sizing the core and bypass ratio of the engine. These studies attempted to fix as few variables as possible for exploring the design space to better assist finding an optimal design. Iterations were used to hold OPR constant, and later used to also hold the ratio between inner fan pressure ratio (IFPR) and outer fan pressure ratio (OFPR) fixed.

5.2.1 Initial Core Scaling (constant fan mass ﬂow) Initial parametric studies held the fan mass ﬂow constant, so varying the bypass ratio scaled the core. Figure 5 shows the result of iterating design bypass ratio and the IPC pressure ratio while holding OPR constant. This was used to increase the bypass ratios for both the 3-spool and geared engine designs. Figure5 also shows that the optimum was more limited in the geared case. The results of this study showed the need to increase the fan size to be able to consider higher bypass ratios while meeting the required thrust.

(a) 3-spooled architecture with max BPR around 12. (b) Geared architecture with max BPR around 11.8.

Figure 5: Initial sizing of the two architectures with BPR and IPC being varied. The contour shows speciﬁc fuel consumption.

5.2.2 Bypass Scaling (constant core scale) The core sizes found from the previous study were held constant while varying the bypass ratio increased the fan. The same parameters were varied and the same iterations were used. This was used to get a better feel for the design space and ﬁnd the compressor pressure ratio balance. These studies showed a clear optimum point where the parametric surface has a maximum. This was used to get higher bypass ratios, and then the full engine would be scaled down so there was not an excess of thrust.

9 (a) 3-spooled architecture. (b) Geared architecture

Figure 6: Optimization of BPR while holding the reduced core size constant. Unlike the previous study, both designs showed potential designs with a design BPR of 14.

5.2.3 Evaluation of Design for Takeoff The parametric studies so far either held the overall mass flow or core mass flow constant. This meant that only one of the fan scale or the core scale were varying at a time. These plots were used to evaluate the converged design by varying both. Previous studies varied the IPC pressure ratio and IFPR, but both were constant for this one. These plots were used to check and refine the results from the optimization for takeoff.

Figure 7: Study varying both fan and core size to check and reﬁne the results of the previous studies. Both plots are for the geared architecture with contours of TSFC on the left and fuel ﬂow on the right.

Previous studies had not considered outer fan pressure ratio (OFPR). It was found that OFPR had a significant effect on the optimum shown in Figure6. The following studies held the ratio of IFPR and OFPR constant. This ratio was set to be 1.02, which was similar to the ratio of values in the baseline. With this change, Figure8 showed clearly different trends observed between the two engine architectures. Up to this point, the two had nearly identical performances, but this study seemed to imply that there was an upper limit to the 3-spool design that was not present in the geared configuration. Note that IFPR, OFPR, IPC pressure ratio, and BPR are all varying in this study due to the implementation of iterations. From this

10 point, the GTF was determined to have higher potential for improvement and chosen as the ﬁnal architecture.

Outer Fan Pressure Ratio = 1 ... 1.8 Outer Fan Pressure Ratio = 1.1 ... 1.8 Design Bypass Ratio = 10 ... 22 Net Thrust < 374[kN] Design Bypass Ratio = 10 ... 22 Net Thrust < 374[kN] Sp. Fuel Consumption [g/(kN*s)] = 5.2...11.2 Sp. Fuel Consumption [g/(kN*s)] = 5.6...11.6 1.7 1.7

1.6 1.6

1.5 1.5

1.4 7 1.4 8 . 2 7 8 6 6. .2 .8 4 7 6 .6 .8 7 4 .6 7 3

8 3 1.3 1.3 . 7 8 4 Outer Fan PressureFan Outer Ratio PressureFan Outer Ratio 6 8 6 .4 .4 5 9 8 .6 . 6 .4 2 9 8. .2 8

1.2 9 1.2 9. .6 6

10 10 1 10 1 0.4 .4 0. 1 8 1 0. 1.2 8 1.1 1.1 8 10 12 14 16 18 20 22 24 8 10 12 14 16 18 20 22 24 Design Bypass Ratio Design Bypass Ratio

(a) 3-spooled architecture. (b) Geared architecture.

Figure 8: Both architectures have essentially the same contours for TSFC, but the boundary of constant thrust shows a clearly diﬀerent trend that allows for higher bypass ratios in the geared design.

5.2.4 Sizing for Cruise Sizing for previous studies sized for takeoff and greatly improved takeoff performance, however the fuel consumption at cruise had not significantly changed from the baseline engine. Since cruise performance was a higher priority for this design, we started optimizing the engine for cruise using the final geared design from the previous studies.

The sizing for cruise used the same parametric studies as before: varying BPR and IPC PR while holding core scale constant (Figure 9a), and then reﬁning the results by varying BPR and core scale (Figure 9b). The green line in Figure 9a shows where the design was moved to when sized for cruise.

Design Bypass Ratio = 10 ... 20 Design Bypass Ratio = 10 ... 20 IP Compressor Pressure Ratio = 8 ... 12 HPC Corr. Flow W25Rstd = 10 ... 15 [kg/s] Net Thrust < 69[kN] Sp. Fuel Consumption [g/(kN*s)] = 13.6...20.4 Ovl_Max Engine Diameter [m] = 3...5.25 12 1.24 20 .6 19 .2 19 .8 11.5 18 4 1.2 15 18.

8 .6 14.7917 1 17 11 14.5833 2 8 1.16 7. 6. 1 1 .4 16 14.375 6 10.5 1 14.1667 2 5. 1.12 1 4 6 8 1 13.9583 6 . 4. .4 9 15 1 14 10 13.75 1.08 13.5417 9.5 13.3333 13.125 FuelFlow [kg/s] 1.04 4 12.9167 .65 9 815 ...9.2 4545 12.7083 4 .5 IP CompressorPressure Ratio 1 12.5 8.5 12.2917 12.0833 8 .96 11.875 11.6667 7.5 .92 8 10 12 14 16 18 20 22 8 10 12 14 16 18 20 Design Bypass Ratio Design Bypass Ratio

(a) Varied BPR and IPC pressure ratio. Green line indi- (b) Varied BPR and core size. Greyed area is below the cates optimal TSFC that was targeted. takeoﬀ thrust limit.

Figure 9: Final parametric studies for sizing the geared engine. These studies were set up the same as previous studies, but sized for cruise.

11 With sizing for cruise, extra steps had to be taken to hold the OPR constant and visualize the minimum thrust requirement at takeoff. OPR at cruise was set to 83, which roughly placed the OPR to 60 at takeoff when evaluated in off-design. The boundary on the parametric study shows a constant thrust line at cruise used to approximate the required thrust at takeoff.

Figure 9b shows a clear optimum for fuel consumption. The off-design team was given design points along the constant thrust line that met the required thrust and OPR at takeoff to evaluate the total mission performance. The mission analysis confirmed that this optimum resulted in the lowest mission total fuel burn. This design point should likely be at a lower engine diameter and bypass ratio, and the following section discusses factors that would have influenced this result.

5.2.5 Further Improvements Upper limitations of engine size were largely unaddressed. The project document did not set constraints for a maximum diameter, but this should have been considered. It would have been good to derive a maximum engine diameter based on the aircraft’s wing geometry relative to the ground.

The team only found that ram drag was included in GasTurb. Skin friction drag should have also been approximated. This could have been done this using composed values to estimate the drag force by roughly calculating outer surface area and assuming a drag coeﬃcient. Then the net thrust minus this estimated drag would have been used in place of net thrust in the previous studies. This would have shifted the minimum thrust boundary depicted in Figure 9b and would have shifted the optimum to a lower point.

Similarly, weight was unaddressed in the off-design mission analysis. There could have been an estimation of the impact of engine weight on the mass fraction of the aircraft, and the corresponding change in required fuel for the mission. This trade-off may have been small since the significant increase in fuel efficiency would reduce the weight of required fuel at takeoff, but it would have been a good to address.

A mixed nacelle design was going to be considered and the file was ready, but time ran out to do a proper comparison with this design change. This would have included changes to the low-pressure turbine such as a possible stage reduction, then refining the core and bypass sizing with the same studies as before to find the optimum with mixing. This also would have needed to include the previously mentioned skin friction drag approximation to be a fair comparison.

5.3 Gearbox As stated in the RFP, the gearbox mass would be calculated based on the torque transmitted to the fan. The gearbox was designed in tandem with the fan and optimized to give ideal performance with respect to thrust, overall engine mass, and fuel consumption. Below are several single-parameter parametric studies that conﬁrm that the operating point (in purple) is very close to the maximum or minimum.

Figure 10: The operating point is also near the point of maximum cruise net thrust.

12 (a) (b)

Figure 11: Parametric studies showing the operating point close to the minimum for both SFC and overall mass.

5.4 Compressors A compressor, by design, increases the pressure along the flow direction. However, fluids want to flow from high to low pressure, which in the compressor is in the negative axial direction. This is what is known as an adverse pressure gradient. In an adverse pressure gradient, there is a possibility that the boundary layer could separate, which can cause stall, leading to compressor surge, where the flow actually reverses and flows out of the front of the engine. This is very bad and must be avoided at all costs.

To avoid boundary layer separation, compression must be done slowly, with small pressure ratios over each stage. As outlined by Boyce [4], the ranges of compressor operation is listed as follows.

Table 5: Compressor characteristics based on application.

Inlet Relative Velocity Pressure Ratio Eﬃciency Sector Flow Regime Mach Number per Stage per Stage Industrial Subsonic 0.4 − 0.8 1.05 − 1.2 88 − 92% Aerospace Transonic 0.7 − 1.1 1.15 − 1.6 80 − 85% Research Supersonic 1.05 − 2.5 1.8 − 2.2 75 − 85%

Consulting the Aerodynamic Design tab within Geometry, we are shown that the booster has a relative rotor 1 inlet Mach number of 0.871. From this, our team determined the range of acceptable compressor stage pressure ratios to be between 1.15 − 1.6. With this, the bounds of pressure ratios (PR) for a component with a known number of stages can be calculated simply using

number of stages PRcomponent = PRstage (1)

5.4.1 LP Compressor (Fan) The fan and LPC characteristics are given in the below table. The LPC tip speed (414.48 m/s), LPC Inlet Radius Ratio (0.283), LPC Inlet Mach number (0.574), and other inputs to the LPC stage were kept constant to the baseline engine. When comparing the output of baseline engine to our redesigned engine it is seen that Design LP Spool speed is reduced from 2471.72 to 2023.02 RPM, while keeping the Corrected ﬂow per unit kg area same (197.18491 s·m2 ). All these values are achieved while increasing the BPR from 9.30 to 17.833. As our key focus was to maintain or improve the Corrected ﬂow from baseline engine, various parametric studies

13 were ran to optimize LPC tip speed to increase the flow rate, but it was realized that varying tip speed does not correlate to Corrected flow. Therefore the following inputs were finalized for LPC design.

Table 6: GasTurb summary of the LPC.

Input: LPC Tip Speed m/s 411.48000 LPC Inlet Radius Ratio 0.28300 LPC Inlet Mach Number 0.57400 Engine Inl/Fan Tip Diam Ratio 1.00000 min LPC Inlet Hub Diameter m 0.00000 Output: LPC Tip circumf. Mach No 1.34528 LPC Tip relative Mach No 1.46262 Design LP Spool Speed [RPM] 2023.02 Design IP Spool Speed [RPM] 3478.67 LPC Inlet Tip Diameter m 3.88462 LPC Inlet Hub Diameter m 1.09935 Calculated LPC Radius Ratio 0.28300 LP Spool Torque N*m 87618.04 Aerodynamic Interface Plane m² 11.85189 Corr.Flow/Area LPC kg/(s*m²) 197.63383

Following finalization of LPC design inputs, off design analysis was conducted to find the optimal operating line. With the initial LPC design the operating line is seen below. At this operating line, the operating point yields a surge margin percentage of 45.43 as seen in the below figure.

1.6 1

1.5

1.4 2 /P

13 1.3 5 .8 0 0 .9 0 2 9 3 . 9 4 0 . .9 0 0 5 .9 1 1.2 0 0.95 0.9 0.8

Pressure Ratio P 1.1

. 0.80 7

0.70 60 0 0. . 6 0.500.40 1 0. 5 0. 0 4 .3

.9 0 250 500 750 1000 1250 1500 1750 2000 2250 Mass Flow W [kg/s] 2RStd Figure 12: Initial operating line for the LPC (fan).

But at the initial operating point the percentage change in Mach number does not equal to percentage change in Corrected flow. As this is not possible, the LPC operating line had to be scaled to a point where both the percentage change in Mach number equaled percentage change in Corrected flow. As seen in the figure below, this optimized line operates closer to the Surge line but still has sufficient margin to not have high concerns for Surge.

14 Figure 13: Optimized operating line for the LPC (fan).

Table 7: Initial operating point. Table 8: Optimized operating point.

Spool Speeds: LP Spool Spool Speeds: LP Spool Absolute [RPM] 2023.0 Absolute [RPM] 2023.0 Relative 1.0000 Relative 1.0000 LPC LPC Surge Margin [%] 45.477 Surge Margin [%] 33.896 Handling Bleed WB,hd/W22 Handling Bleed WB,hd/W22

The final LPC design yields an outer LPC polytropic efficiency of 0.9147 with an inner LPC polytropic efficiency of 0.8951. All this is achieved while almost doubling the BPR compared to the baseline engine.

5.4.2 IP Compressor (Booster) With 8 stages, the intermediate pressure compressor can have a pressure ratio ranging from 1.158 to 1.68, or between 3 and 43.

One thing we noticed while doing oﬀ-design analyses was that the bottom part of the operating line for the booster meandered past the surge line. This is not acceptable for performance, as discussed earlier, the engine will stop working correctly. So in order to maintain the safety and integrity of the engine, this had to be corrected. To address this, a handling bleed was used. As explained in the GasTurb 13 [5] user manual, the handling bleed discharges some of the compressed air into the bypass duct or overboard, which helps to lower the operating line of the compressor and avoid a surge. While a static bleed fraction of the mass ﬂow can be entered into the secondary inputs window, GasTurb also allows a Handling Bleed Schedule, which automatically varies the amount of bleed air based on the engine RPM.

15 (a) IPC without handling bleed scheduling (b) IPC with handling bleed scheduling.

Figure 14: Compressor maps of the booster (IPC) at cruise conditions with and without a handling bleed.

As shown above in Figure 14a, without the handling bleed, the operating line passes the surge line, which is unacceptable. Once a scheduled handling bleed is employed, the entire operating line (at cruise conditions) stays below the surge line.

5.5 LP Turbine The design of LP turbine is critical and challenging, LP turbine stage is one of the major component that consumes lot of weight in the overall engine weight. The critical design of this LP turbine is based on decreasing the weight of the component and doing so by increasing performance of the baseline LP turbine. The baseline LP turbine is a 7 stage axial-ﬂow turbine with some of the key parameters as shown below.

Table 9: Design output from GasTurb for baseline LPT

Length m 0.747555 Total Number of Blade and Vanes 5009 Casing Mass kg 87.7429 Total Vane Mass kg 158.865 Total Blade Mass kg 288.567 Inner Air Seal Mass kg 26.2607 Rotating Mass kg 595.208 Total Mass kg 841.817 Polar Moment of Inertia kg*m2 307.242

GasTurb 13 allows us to modify 3 crucial parameters namely: • LP Exit Diameter • LP Inlet Diameter • LP Exit Radius ratio

These parameters influence the overall LP turbine characteristics and optimizing these parameters is the key so as to reduce weight of the engine, specific fuel consumption and the mass flow through the engine which is our primary objective in the current project.

16 5.5.1 Optimization of Speciﬁc Fuel Consumption

Legend:

Figure 15: Parametric Study: LPT Exit Dia (vs Inlet dia & vs exit radius ratio) for range of SFC.

g The black point in the graph is the optimized point with specific fuel consumption of 14.8 kN·s . One can observe better value of specific fuel consumption (SFC) in the graph but as seen in the graph if the engine is modified to that dimension the inlet diameter is increased by 0.2 m which increases the weight of the component by 50 kgs. Radius ratio is increased to 0.85 which is high value and not recommended by the GasTurb software and can having degrading effect to engine performance. These changes are not feasible as g it decreases the SFC by 0.1 kN·s .

5.5.2 Optimization of Core Flow The core flow is not influenced greatly by the dimensions and performance of the LP turbine. The graph below shows the core flow is at 12.2 kg/s which is significantly optimized compared to 14.1 kg/s fuel flow in the baseline engine. Further decreases of this value again demands larger change in the diameters and hence affecting weight of the engine.

Figure 16: Parametric Study: LPT exit diameter vs core ﬂow for minimum SFC.

17 5.5.3 Final Results for LP Turbine Table 10 below shows the changes in the new turbine compared to the baseline engine, as we can see there is a signiﬁcant improvement in terms of weight. The blades and vanes are increased by a fractional margin and there is also a slight increase in the blade height which might increase the cost of production of these blades but the overall length of the LP turbine is decreased by 40%.

The LP Turbine velocity triangles for the new engine are shown below in Figure 17. As we can see, the ﬂuid ﬂow at exit of the last stage is more axial which means the tangential component of the exit velocity is less and hence there is better power extraction by the LP turbine.

Table 10: Design output from GasTurb for new LPT

Length m 0.440371 Total Number of Blade and Vanes 5197 Casing Mass kg 40.5966 Total Vane Mass kg 62.8789 Total Blade Mass kg 125.236 Inner Air Seal Mass kg 8.22844 Rotating Mass kg 538.701 Total Mass kg 642.176 Polar Moment of Inertia kg·m2 130.319

Figure 17: LPT velocity triangles.

The eﬃciency is slightly better than baseline engine this is also another key improvement in the engine and the stag loading is also slightly decreased. The overall performance data is shown below in Table 11.

18 Table 11: LPT detailed performance of the new LP turbine.

Input: Number of Stages 7 LPT with EGV's [0/1] 1.00000 1. LPT Rotor Inlet Dia m 1.50000 Last LPT Rotor Exit Dia m 2.15000 LPT Exit Radius Ratio 0.80000 LPT Vax.exit / Vax.average 1.00000 LPT Loss Factor [0.3...0.4] 0.35000 LPT 1. Rotor Cooling Constant 0.00000 Output: LPT Inlet Radius Ratio 0.95521 LPT First Stator Exit Angle 65.65519 LPT Exit Mach Number 0.35556 LPT Exit Angle -5.90252 LPT Last Rotor abs Inl Temp K 887.52188 LPT First Rotor rel Inl Temp K 1427.00 LPT First Stage H/T J/(kg*K) 84.22955 LPT First Stage Loading 1.64847 LPT First Stage Vax/u 0.47466 LPT Exit Tip Speed m/s 435.11900 LPT Exit A*N*N m²*RPM²*E-6 19.52591 LPT 1.Rotor Cool. Effectiveness 0.00000 LPT 1.Rotor Bld Metal Temp K 1427.00 LPT Torque N*m 87618.04

Velocities: 1st Stage Inlet Absolute Velocity V m/s 314.59 1st Stage Inlet Axial Velocity Vax m/s 129.68 1st Stage Inlet Relative Velocity W m/s 130.38 1st Circumferential Velocity U m/s 273.21 1st Stage Exit Absolute Velocity V m/s 130.38 1st Stage Exit Axial Velocity Vax m/s 129.68 1st Stage Exit Relative Velocity W m/s 314.59

Last Stage Inlet Absolute Velocity V m/s 450.92 Last Stage Inlet Axial Velocity Vax m/s 185.88 Last Stage Inlet Relative Velocity W m/s 186.87 Last Circumferential Velocity U m/s 391.61 Last Stage Exit Absolute Velocity V m/s 186.87 Last Stage Exit Axial Velocity Vax m/s 185.88 Last Stage Exit Relative Velocity W m/s 450.92

5.6 Materials One risk that was identified early was material degradation and failure due to the higher turbine inlet temperature design requirement of 1930 K. Based on the density of the material used in the baseline engine, it was determined that the turbine disks were a type of nickel-based super-alloy (possibly Rene 41). This is adequate for the 1784 K inlet turbine temperature of the baseline engine but not for the higher temperature the customer requires. A chemical vapor infiltration carbon/silicon (CVI-C/Si) ceramic matrix composite (CMC) material was chosen to replace the nickel-based super-alloy. The CVI-C/Si material has a lower thermal expansion coefficient and a higher specific heat which allows it to handle higher temperatures without effecting property characteristics. Property characteristics of Rene 41 beyond 1000 K were not able to be obtained, however, based on the below table it can be inferred that strength properties between the two materials at higher temperatures would be similar meaning that the structural integrity of the turbine disks will not be affected by switching material to this new CMC material.

19 Table 12: Properties for Turbine Disk Materials (*Properties taken at 1000 K). [1], [5].

CVI-SiC/SiC Rene 41* Density (kg/m3) 2100 8250 Ultimate Tensile Strength (MPa) 310 304 Elongation % 0.75 20 Modulus of Elasticity (GPa) 95 149 Thermal Expansion (E-6/K) 2.3 16.9 J Speciﬁc Heat kg·K 1134 452 W Thermal Conductivity m·K 10 15.8 Yield Tensile Strength (MPa) 475 345 Operating Temperature (K) > 2000 < 1800

Material degradation and failure in the compressor was also a concern since our design has higher compressor pressure ratios than the base line engine. These higher ratios would lead to increased stresses being experi- enced in the compressor. Based on the density of the material used in the base line engine, it was determined that the compressor was another type of a nickel-based super-alloy (possibly from the Inconel family). To handle the increased stress while also attempting to decrease overall weight, an aluminum-iron (ntAl-xFe) material that is currently being researched and developed at Purdue was chosen [10], [6]. There is not a lot of information available for this material, however, initial research shows that it has the potential to be stronger and more ductile than stainless steels and even nickel-based alloys at about 30%-50% of the weight.

Figure 18: Speciﬁc Strength vs. Speciﬁc Modulus for Alloys including nt Al-xFe [10].

Strength versus temperature plots for this new aluminum-iron material could not be acquired so for the later stages of the compressor a titanium-alloy material was chosen. The temperature proﬁle of the compressor fell within the range where speciﬁc strength for the titanium alloy was higher than that of the nickel alloy as seen in the below graph.

20 Figure 19: Speciﬁc Strength vs. Temperature for Various Alloys [12].

In addition to this beneﬁt, the density of this new material is signiﬁcantly lower than that of the nickel-based alloy which resulted in lower overall weight of the engine.

As mentioned above, the nickel-based super-alloy material in the turbine components was replaced with a carbon/silicon CMC material that resulted in our engine being able to handle the required higher turbine inlet temperature while simultaneously lowering the overall weight. Also mentioned above was the decision to replace the material in the high-pressure compressor with the ntAl-xFe and titanium alloy materials. These materials were better able to handle the increased stresses seen in the compressor while also reducing overall weight. Because of the aluminum-iron material’s low density and high strength, it was also implemented into the compressor casing, interduct casing, low pressure turbine casing and disks, and exhaust casing. The remainder of the materials were left similar to the baseline engine since changes to these would not result in lower weights or more favorable mechanical and/or thermal properties.

5.7 Engine Cycle Improvements Improvements to the overall engine cycle were explored in order to fulfill the customer’s requirement of increasing engine efficiency over the mission of the aircraft. An afterburner (or a reheat) was briefly discussed but immediately dismissed as this would significantly increase thrust (not a customer requirement) at a cost of increasing fuel consumption and decreasing efficiency. As mentioned earlier, a mixed flow design was considered as well however this would result in an increased geometry design for the nacelle. Initial mixed vs. unmixed cycle analysis was run yielding comparable results. Because of this, the mixed option was dismissed. An intercooler was considered as well since this would result in lower compression work, which would lead to the expansion requirement of the turbine decreasing, which would ultimately result in a lower turbine inlet temperature. In addition, regeneration (or a heat exchanger) was explored simultaneously since this would allow the flow at the exit of the compressor to be heated up without burning and fuel resulting in higher engine efficiency. Effectiveness of the intercooler and heat exchanger were taken as 0.6 and 0.7 respectively as they are realistic values [11]. A trade-off study was performed between implementing an intercooler or a heat exchanger or a combination of the two and compared to the baseline engine with the below results.

Table 13: Intake Mass Flow: 1561.0 kg/s. Tradeoﬀ Study with Intercooler and Heat Exchanger

Base Engine w/ Base Engine w/ Base Engine Base Engine Intercooler Only Intercooler & HE w/ HE only FN (kN) 407.18 407.05 401.63 369.45 g TSFC kN·s 6.6953 7.6768 7.1478 6.6876 WF (kg/s) 2.72617 3.1249 2.8708 2.4707 Core Eﬀ. 0.5185 0.5075 0.5468 0.523 s Nox 3.5924 0.9536 1.6342 3.7506

21 This study shows that decreasing the intake mass flow further increased TSFC while further decreasing core efficiency. From these two studies it was determined that adding only a heat exchanger yielded the best results based on the customer’s requirements. A summary of these results, as well as overall weights of both configurations after new material implementation, is seen in the below table.

Table 14: Results of Tradeoﬀ Study of Cycle Improvements (Original Base Engine Weight:7373.4 kg).

Base Engine w/ Base Engine w/ New Materials Difference % Difference New Materials & Heat Exchanger Thrust (kN) 407.18 403.94 -3.24 -0.80 g TSFC kN·s 6.6953 6.673 -0.0223 -0.33 WF (kg/s) 2.72617 2.6955 -0.03067 -1.13 Core Eff. 0.5185 0.5252 0.0067 1.29 Overall Weight (kg) 6997.85 7074.8 76.95 1.10

TSFC and WF were decreased while core eﬃciency was increased all while keeping the overall thrust within the customer’s requirements. With the new materials, the addition of a heat exchanger only adds ≈ 77 kg to the overall weight of the engine. With the original base engine weighing in at 7373.4 kg, the combination of the new materials and heat exchanger still resulted in a 4.05% decrease in overall weight. Thermodynamic cycles, along with other properties, of the base line engine and our new design were taken from GasTurb and are seen below. The outlet of the compressor was able to be increased with the implementation of a heat exchanger as expected.

5.8 Risks and Concerns Of general concern for any engineering task is economics and mechanical feasibility and compatibility. Design for manufacturing (DFM) and design for assembly (DFA) techniques should be analyzed in order to ensure the ﬁnal design is comparable to the baseline engine with regards to process capability, installation, service, and overhaul procedures. Cost analysis should also be performed in order to ensure the budget is within the customer’s requirements.

Drag One concern was the higher engine bypass ratio requirement. This requirement results in a larger fan pressure ratio which necessitates a larger diameter fan. As a result, engine size increases along with overall weight and drag. With the lower density materials listed in Table 37, overall weight was decreased by 5%. This decrease in overall weight leads to a decrease in the induced drag of the engine.

Recall from an introductory aerodynamics course that induced drag (in level ﬂight) can be expressed as

W 2 D = i 1 2 2 2 ρV πb where W is the weight of the aircraft, ρ is the ambient air density, V is the ﬂight velocity, is the wing’s Oswald eﬃciency factor, and b is the wingspan.

Because the induced drag is a function of the square of the weight, it is believed that this decrease in overall weight will overcome the increased geometry (and therefore greater pressure drag) of our design.

Noise Pollution Noise pollution was also a concern that was addressed. One customer requirement that works in the favor of this issue is the higher bypass ratio. This requirement means that more air is forced around the engine

22 core. The resulting engine design has a larger, more slowly-rotating fan with fewer blades which are features that reduce an engine’s noise profile. However, as mentioned before, a higher bypass ratio leads to higher compressor pressures ratios which result in higher mechanical speeds. From GasTurb, the base engine requires a high-pressure compressor (HPC) speed of approximately 12,000 RPM while the new design requires a HPC speed upwards of 14,500 RPM. It was decided to implement a gearbox into the design which suppresses engine noise by decoupling the fan and allowing it to rotate at a slower speed. It is difficult to quantitatively estimate the reduction in noise, however, Pratt & Whitney claim that their GTF model engine results in a 75% reduction in noise footprint [9]. Because of this it is believed that our gear reducer will make up for the 20% increase in the HPC speed thus resulting in no significant increase in noise. A mixed flow design was also explored as an option to reduce noise. However, initial analysis showed no significant increase in engine performance, so it was decided that the tradeoff of increasing the overall geometry of the engine to establish a mixed flow was not feasible. Other options to explore in the future to reduce noise include implementing blended wing designs, morphing geometry, and acoustic liners.

6 Our Engine Solution

Table 15: Comparison summary of our team’s ﬁnal engine design compared to the baseline engine.

Metric Baseline Our Design Change Total Mass [kg] 6963 5615 -19.36% g SFC (cruise) [ kN·s ] 18.1933 14.3785 -20.97% m˙ fuel (cruise) [kg/s] 1.05338 0.94863 -9.94% Propulsive Eﬃciency [%] 81.13 83.83 +3.33% BPR 9.3 17.8333 +91.76% Thrust (cruise) [kN] 57.90 65.98 +13.96% Thrust (takeoﬀ) [kN] 396 374.33 -5.48% Turbine Inlet Temperature T4 [K] 1784 1930 +8.2% Overall Pressure Ratio p3/p2 50 60 +20%

Figure 20: Final engine model schematic, geared 2-spool.

23 Figure 21: Thermodynamic cycle of the ﬁnal engine with intercooling and regeneration.

7 Oﬀ-Design Analysis

Off-design studies are related to the performance/behavior of the engine with given geometry at different operating conditions. The geometry is found by running a single cycle design point. Off-design analysis gives us unique insight into the engine design in terms of SFC, fuel flow, and thrust at different conditions other than the cruise conditions at which the design is initially optimized for.

After the two engine files – baseline engine (3 spool ungeared) and the final engine (2 spool geared), were finalized by on-design performance optimization, the off-design tab of GasTurb was used to compare the performance of both the engines with respect to each other. The major performance parameters considered were: time-averaged SFC, time-averaged fuel flow and total fuel burn. To calculate these quantities, the entire flight/mission profile for the engines had to be defined from the flight design points in the RFP. The design specified in the RFP were as follows:

• Takeoﬀ at sea-level

• Cruise at an altitude of 12,190 m at Mach No of 0.85

• Range of flight must be 17,600 km The following sections describe the procedure followed to define the mission profile, calculate the performance of the engines at different mission legs and finally present the comparison of results between the engines.

7.1 Off-Design Operating Point The first step of off-design analysis is to calculate the off-design operating point for the engines to make sure that the results are reasonable and the values are within acceptable ranges. The operating point was calculated at sea level and cruise to confirm the expected values. The operating point for cruise is included below as an example.

24 Table 16: Summary of the ﬁnal engine at cruise conditions.

W T P WRstd Station kg/s K kPa kg/s FN = 65.98 kN amb 216.65 18.760 TSFC = 14.3785 g/(kN*s) 2 689.863 248.02 30.097 2154.741 WF = 0.94863 kg/s 13 653.233 278.73 43.710 1489.329 s NOX = 1.9781 21 36.630 277.68 42.853 85.023 22 36.630 277.68 42.853 85.023 Core Eff = 0.6010 24 36.630 563.80 432.026 12.017 Prop Eff = 0.8383 25 36.630 563.80 425.545 12.200 BPR = 17.8333 3 34.798 950.80 2446.885 2.618 P2/P1 = 1.0000 31 30.403 950.80 2446.885 P3/P2 = 81.30 4 31.351 1930.00 2349.010 3.500 P5/P2 = 0.9482 41 33.183 1880.88 2349.010 3.657 P16/P13 = 0.9750 43 33.183 1522.23 750.006 P16/P6 = 1.50836 44 35.381 1489.52 750.006 P16/P2 = 1.41599 45 37.055 1460.91 746.600 11.323 P6/P5 = 0.99000 49 37.055 745.57 28.539 A8 = 0.94659 m² 5 37.212 745.16 28.539 212.458 A18 = 6.70135 m² 8 37.579 747.13 28.254 217.003 XM8 = 0.79967 18 649.967 278.73 42.617 1519 .879 XM18 = 1.00000 Bleed 0.000 950.80 2446.884 WBld/W2 = 0.00000 ------CD8 = 1.00000 Efficiency isentr polytr RNI P/P CD18 = 0.94000 Outer LPC 0.9101 0.9147 0.355 1.452 PWX = 200.0 kW Inner LPC 0.8897 0.8951 0.355 1.424 V18/V8,id= 0.83750 IP Compressor 0.8929 0.9210 0.442 10.082 WBLD/W22 = 0.00000 HP Compressor 0.8587 0.8864 1.888 5.750 Wreci/W25= 0.00000 Burner 0.9995 0.960 Loading = 100.00 % HP Turbine 0.8541 0.8373 2.590 3.132 WCHN/W25 = 0.05000 LP Turbine 0.9024 0.8587 1.104 26.161 WCHR/W25 = 0.06000 ------WCLN/W25 = 0.04571 HP Spool mech Eff 0.9900 Nom Spd 16308 rpm WCLR/W25 = 0.00429 LP Spool mech Eff 0.9990 Nom Spd 2023 rpm WBLD/W25 = 0.00000 IPC & LPT Nom Spd 3479 rpm Gear Rat = 1.71954 ------WLkBy/W25= 0.00000 P22/P21=1.0000 P25/P24=0.9850 P45/P44=0.9955 WlkLP/W25= 0.01000 ------hum [%] war0 FHV Fuel 0.0 0.00000 43.124 Generic

7.2 Mission Profile The next step involved defining a mission profile for the engines. This can be used to calculate the time and range for each mission leg. Reasonable assumptions had to be made to find the mission profile. To start off, the climb/descent rate and the equivalent air speed for a typical Boeing 787 flight was found.

Table 17: Climb/descent rates and air speed for a typical Boeing 787 [3].

Mission leg Altitude (ft) Climb/descent rate (ft/min) Air Speed (knots) Climb 1 0-18,000 2500 250 Climb 2 18,000 - 26,000 1800 300 Climb 3 26,0000 - cruise 1000 Mach No Descent 1 Cruise - 10,000 1700 250 Descent 2 10,000 - 0 700 140

From the climb/descent rates and air speed, time for each mission leg except cruise was calculated. These times were used to calculate the ranges for respective mission legs except the cruise. With prior information of range being 17,600 km, the range for cruise was calculated by subtracting sum of ranges of other mission legs from 17,600 km. From this, the time for cruise was calculated.

25 Table 18: Flight mission details.

Climb 1 Climb 2 Climb 3 Cruise Descent 1 Descent 2 Altitude (m) 5,486 7,925 12190 12190 3048 0 Airspeed (m/s) 128.61 154.33 250.83 250.83 128.61 72.02 Time (s) 432 266.4 979.32 68231.38 1140.6 857.4 Range (m) 56550.31 47605.15 245647.7 17,114,818 97448.3 37930.52

The TOF (time of flight) obtained was approximately 19.97 hours. This is comparable to a typical New York–Sydney flight which has a range of about 18,000 km. The flight profile is presented in the graph below.

14000

12000

10000

8000

6000

4000

2000

0 0 200 400 600 800 1000 1200

Figure 22: Flight proﬁle for a 17,600 km 787 ﬂight.

7.3 Flight Envelope Once the mission points are defined, the information can used to find values of SFC, fuel flow and thrust at the mission altitudes using the Mission option of GasTurb. For this, we need information about the variation of Mach No. with the altitude for our mission. This can be calculating the flight envelope for the given design points in GasTurb. The following settings were used in flight envelope of GasTurb:

• Turbine Entry Temperature max limiter = 1930 K

• Maximum altitude = 12,190 m

• No. of Altitude Steps = 13

• Minimum Air Speed = 130 knots

• Maximum Air Speed = 300 knots

• Maximum Mach No. = 0.85

• No. of Mach No. steps = 13

Using the above options, the ﬂight envelope obtained is shown below.

26 Figure 23: Operational ﬂight envelope from GasTurb.

The ﬂight envelope gives diﬀerent values of Mach No. for a given altitude depending on air speed. So, a simple Matlab script was written to choose a reasonable variation of Mach No. with Altitude for our mission analysis. The plot of Mach No. vs Altitude (m) from Matlab is shown as follows:

0.9

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

0 0 2000 4000 6000 8000 10000 12000 14000

Figure 24: Mach vs Altitude (m) proﬁle for the 17.600 km 787 ﬂight.

27 7.4 Mission Points With the flight profile and mission points defined, the Mission option was used to find the output (Mainly SFC, fuel flow and thrust) for the series of operating conditions. The detailed output of the mission points from GasTurb is shown below.

Table 19: Results from GasTurb at diﬀerent mission points.

Parameter Units Climb 1 Climb 2 Climb 3 Cruise Descent 1 Descent 2 Altitude m 0 5486 7925 12190 3048 0 Mach Number 0 0.411 0.58 0.85 0.26 0.13 Power Oﬀtake kW 200 200 200 200 200 200 Net Thrust kN 374.335 157.398 114.608 65.9759 215.974 311.451 Core Nozzle Gross Thrust kN 21.3111 18.0469 17.8609 15.3492 19.1757 21.5515 Bypass Nozzle Gross Thrust kN 353.024 293.767 277.131 223.716 314.053 365.447 Gross Thrust kN 374.335 311.814 294.992 239.065 333.228 386.999 g Sp. Fuel Consumption kN·s 6.49903 10.3238 12.0461 14.3784 8.97104 7.84535 Speciﬁc Thrust m/s 223.242 133.476 113.672 95.6363 157.283 182.379 Fuel Flow kg/s 2.43282 1.62495 1.38058 0.948631 1.93751 2.44345 Overall Pressure Ratio p3/p2 60.3027 72.4633 78.172 81.3 66.9484 59.977

With the values at different mission points, a Matlab script was written to find the time-averaged values of SFC, fuel flow and total fuel burn was calculated. The average value of SFC and fuel flow between two mission points was used as the value for that mission leg. This is an approximate assumption which will provide us results very close to that of an actual flight. The following formulas were used to find the time-averaged values.

5 1 X SFC = SFC · ∆t avg T i i i=1 5 1 X WF = WF · ∆t avg T i i i=1 where SFCi and WFi are values for a given mission leg, ∆ti is the time for the given mission leg, and T is the total time of ﬂight.

7.5 Results The results for the final engine design is displayed in two tables. The first table presents how the engine performs through the different legs of the mission and the second table presents improvements from the baseline engine. This is done as a proof of concept.

The following table is a breakdown of the engines operation as it advances through the mission. The total time spent in each leg of the trip is listed alongside fuel flow and fuel burned during each section of the mission. The majority of fuel burned, as expected, is during the cruise condition. But, it should also be noted that during cruise, the lowest fuel flow rate is achieved confirming the optimization process.

28 Table 20: Fuel burn for each mission leg for ﬁnal engine design.

Takeoﬀ Climb 1 Climb 2 Cruise Descent 1 Descent 2 Time (s) 432 266 979 68231 1141 857 Fuel Flow (kg/s) 2.43 1.62 1.38 0.94 1.93 2.44 Fuel Burn (kg) 1,050 431.6 1,351.5 64,137.5 2,201.4 2,092.1

The following table of results summarizes the overall mission performance of both the baseline engine and the final engine. Savings in time weighted SFC and fuel flow are also included. A savings of 21.04 percent was seen for SFC, and 11.51 percent savings for both fuel flow and total fuel burn. A substantial improvement was made from the baseline engine to the final configuration.

Table 21: Comparison of average SFC, fuel ﬂow and total fuel burn.

Parameter Baseline Final Change Time-averaged SFC (g/(kNs)) 17.8578 14.1005 -21.04 % Time-averaged Fuel Flow (kg/s) 1.0810 0.9566 -11.51 % Total Fuel Burn (kg) 77,731 68,789 -11.51 %

The following plots show the comparison of SFC, Thrust and fuel flow for both the baseline (blue) and final (red) engine. Lower thrust is required, and lower fuel flow is observed for the final engine for all the legs of the mission. The SFC is also significantly lower for each leg of the mission.

29 400

300

FINAL 200 BASELINE

100

0 0 200 400 600 800 1000 1200

FINAL BASELINE 10

5 0 200 400 600 800 1000 1200

3.5

2.5

FINAL 2 BASELINE

1.5

0.5 0 200 400 600 800 1000 1200

Figure 25: Variation of Thrust (a) , SFC (b) and fuel ﬂow (c) for the entire mission.

8 Trade-oﬀ Study

Given in the RFP were some design constraints and major design parameters that directly impact or limit the capabilities of the engine. The three most important input parameters were determined to be: turbine entry temperature, bypass ratio, and fan pressure ratio. These parameters are outlined in Table 22. Bypass ratio plays a large role in the project because it is the basis for the entire study of an ultra-high bypass ratio turbofan so it will closely watched in the parametric studies.

30 The output parameters were chosen as they were deemed the most important for a passenger jet engine configuration. Overall engine mass, engine size, core efficiency, and fuel burn will all be scrutinized in the parametric studies. It should be noted that engine size was defined mathematically within the GasTurb program as overall diameter multiplied by overall length so that size could be quantified. Also, fuel burn was defined mathematically as well using thrust specific fuel consumption multiplied by net thrust to give units of grams per second.

Table 22: Trade-oﬀ study parameters

Input Parameter Output Parameter Turbine Entry Temperature Overall Mass Design Bypass Ratio Engine Size Inner Fan Pressure Ratio Core Eﬀ. - Fuel Burn

8.1 Parametric Study The plots that follow in the next section will be structured as follows:

• BPR vs IFPR vs (Ov. Mass, Fuel Burn, SFC)

• BPR vs TET vs (Eng. Size, SFC, Core Eﬀ.) As noted earlier, bypass ratio (BPR) is a major design point so it will be included in both plots to better track the behavior of the engine as bypass ratio is increased in order to determine the best performance trade oﬀs.

8.2 Baseline Engine First, the baseline engine will be modelled and plots will be made using the above parameters and parametric combinations. The operating point will be highlighted with the color purple and a summary of the output will also be included below.

Figure 26: Baseline engine; BPR vs IFPR vs (Ov. Mass, Fuel Burn, SFC).

31 Figure 27: BPR vs TET vs (Eng. Size, SFC, Core Eﬀ.)

In the first figure it can be seen that there is definitely room for improvement and room for the bypass ratio to increase so that SFC may also move to a more favorable region. Also, increasing the BPR will also help move fuel burn to a lower value as well according to the parametric plot.

In the second ﬁgure it can be seen that increasing the BPR will lead to a lower SFC as observed before. It can also be deduced from the plot that increasing burner exit temperature can lead to higher core eﬃciency as well.

The baseline model was presented in order to give a better understanding of the direction of the redesign of the engine and also to give reference to the improvements made in the ﬁnal engine design.

8.3 Geared Solution

Figure 28: Geared Engine; BPR vs IFPR vs (Ov. Mass, Fuel Burn, SFC).

32 Figure 29: Geared engine; BPR vs TET vs (Eng. Size, SFC, Core Eﬀ.)

Looking at the first plot it can be seen that increasing or decreasing the FPR will not have a large effect on SFC as it mainly impacts fuel flow. Increasing the BPR trends in the direction of lower SFC but will increase the overall engine mass possibly negating the effects of a lower SFC. For the reason of increasing overall mass with BPR, efficiency’s have to be taken into consideration.

The second plot elaborates on the effect of BPR and efficiencies needed to further understand what trade offs need to be made. A high BPR trends to a lower SFC as before but it may also now be seen that core efficiency increases as well. A higher TET coupled with a higher BPR leads to a higher core efficiency. The only negative impact observed from increasing the BPR is overall engine size increases as well which can lead to higher engine mass. The trade off made was to increase the BPR, TET, and use different materials in the engine to reduce the weight penalty.

9 Limits of Analysis

Atmospheric Conditions Throughout the entirety of this project, our analysis was done using dry air and standard atmosphere conditions. This is a huge deviation from reality, at least at takeoﬀ. Relative humidity values can easily reach over 80% in Indiana.

Furthermore, standard sea-level conditions are deﬁned as 15◦ C. Since pressures and densities (and therefore thrust) vary quite a bit with temperature, a change in ambient temperature at an airport could drastically change the engine’s performance, so much to the point that some aircraft might not be able to take oﬀ with the given runway length available.

33 10 Appendix A - Velocity Triangles

Booster

Figure 30: Velocity Triangles for the Booster (IPC).

HP Compressor

Stage 1 Stage 2 Stage 3

a = 18.8° b = -49.6° a = 17.8° b = -47.9° a = 17.8° b = -47.9° Mv = 0.504 Mw = 0.736 Mv = 0.529 Mw = 0.751 Mv = 0.516 Mw = 0.734

Mu = 0.722 Mu = 0.719 Mu = 0.702

a = 49.6° b = -18.8° a = 47.9° b = -17.8° a = 47.9° b = -17.8° Mv = 0.718 Mw = 0.491 Mv = 0.733 Mw = 0.516 Mv = 0.717 Mw = 0.504

Mu = 0.704 Mu = 0.701 Mu = 0.686

Stage 4 Stage 5 Stage 6

a = 17.0° b = -48.3° a = 17.8° b = -47.9° a = 17.8° b = -47.9° Mv = 0.471 Mw = 0.678 Mv = 0.502 Mw = 0.714 Mv = 0.488 Mw = 0.693

Mu = 0.663 Mu = 0.644 Mu = 0.683

a = 47.9° b = -17.8° a = 46.7° b = -20.2° a = 47.9° b = -17.8° Mv = 0.679 Mw = 0.478 Mv = 0.645 Mw = 0.472 Mv = 0.699 Mw = 0.491

Mu = 0.650 Mu = 0.632 Mu = 0.669

Figure 31: Velocity Triangles for the HPC.

34 11 Appendix B - GasTurb Engine Conﬁguration (Final GTF)

Table 23: GasTurb inputs to Basic Data.

Property Unit Value Comment Intake Pressure Ratio 1 No (0) or Average (1) Core dP/P 1 Inner Fan Pressure Ratio 1.42382 Outer Fan Pressure Ratio 1.45229 Core Inlet Duct Pressure Ratio 1 IP Compressor Pressure Ratio 10.0817 Compressor Interduct Pressure Ratio 0.985 HP Compressor Pressure Ratio 5.75 Bypass Duct Pressure Ratio 0.975 Turb. Interd. Ref. Press. Ratio 0.98 Design Bypass Ratio 17.8333 Burner Exit Temperature K 1930 Burner Design Efficiency 0.9995 Burner Partload Constant 1.6 Used for off design only Fuel Heating Value MJ/kg 43.124 Overboard Bleed kg/s 0 Power Offtake kW 200 HP Spool Mechanical Efficiency 0.99 Gear Ratio 1.71954 LP Spool Mechanical Efficiency 0.999 Burner Pressure Ratio 0.96 Turbine Exit Duct Press Ratio 0.99

Table 24: GasTurb inputs to Secondary Data.

Property Unit Value Comment Rel. Handling Bleed to Bypass 0 Rel. HP Leakage to Bypass 0 Rel. Overboard Bleed W_Bld/W25 0 Rel. Enthalpy of Overb. Bleed 1 Recirculating Bleed W_reci/W25 0 Rel. Enthalpy of Recirc Bleed 1 Number of HP Turbine Stages 1 HPT NGV 1 Cooling Air / W25 0.05 HPT Rotor 1 Cooling Air / W25 0.06 HPT Cooling Air Pumping Dia m 0 Calculated in HPT Design Number of LP Turbine Stages 7 LPT NGV 1 Cooling Air / W25 0.0225 LPT Rotor 1 Cooling Air / W25 0.0025 LPT NGV 2 Cooling Air / W25 0.0225 LPT Rotor 2 Cooling Air / W25 0.0025 Rel. Enth. LPT NGV Cooling Air 0.6 Rel. Enth. of LPT Cooling Air 0.2 Rel. HP Leakage to LPT exit 0.01 Rel. Fan Overb.Bleed W_Bld/W13 0.005 Core-Byp Heat Transf Effectiven 0 Coolg Air Cooling Effectiveness 0 Bleed Air Cooling Effectiveness 0

35 Table 25: GasTurb inputs to the Engine Inlet tab under Geometry.

Number of Struts 0 Strut Chord/Height 0 Gap Width/Height 0.2 Cone Length/Radius 2.2 Cone Angle deg 28.5 Casing Length/Radius 0.3 Casing Thickness m 0.005 Casing Material Density kg/m3 1601.85 Inlet Mass Factor 1

Table 26: GasTurb inputs to the Fan tab under Geometry.

Number of Stages 1 Inlet Guide Vanes (IGV) 0/1 0 IGV Profile Thickness % 5 IGV Material Density 1281.48 Annulus Shape Descriptor 0…1 0.655 Inlet Radius Ratio 0.3 First Stage Rotor Aspect Ratio 3 Last Stage Rotor Aspect Ratio 1.5 Core Vane Aspect Ratio Span/Chord 1.908 Bypass Vane Aspect Ratio 2.75 Core Vane Gap/Chord Ratio 0.84 Bypass Gap/Chord Ratio 1.02 Rotor Pitch/Chord Ratio 0.708 Core Vane Pitch/Chord Ratio 0.4 Core Exit Vane Gap/Chord Ratio 0 Core Exit Duct Radius Ratio 0.6034 Bypass Vane Pitch/Chord Ratio 0.15 Bypass Vane Lean Angle 8 Bypass Inner/Splitter Radius 1 Disk Bore / Inner Inlet Radius 0.05 Rel Thickness Inner Air Seal 0.04 Casing Thickness 0.005 Casing Material Density kg/m3 1601.85 Containment Ring Thickness % 5 Containment Ring Mat Density kg/m3 320.369 Mean Bypass Vane Thickness % 5 Byp Vane Material Density kg/m3 1601.85 LP Compressor Mass Factor 1

36 Table 27: GasTurb inputs to the Booster tab under Geometry.

Number of Stages 8 Number of Variable Guide Vanes 0.4 Inlet Guide Vanes (IGV) 0/1 1 IGV Profile Thickness % 5 IGV Material Density kg/m3 1601.85 Annulus Shape Descriptor -0.5…1 -0.1 Blade and Vane Sweep 0/1 0 First Stage Aspect Ratio 2.776 Last Stage Aspect Ratio 2.412 Blade Gapping: Gap/Chord 0.12 Pitch/Chord Ratio 0.7 Disk Bore / Inner Inlet Radius 0.3 Rel Thickness Inner Air Seal 0.04 IP Compressor Mass Factor 1 Casing Thickness m 0.005 Casing Material Density kg/m3 1601.85 Casing Thermal Exp Coeff E-6/K 10 Casing Specific Heat J/(kg*K) 500 Casing Time Constant 10 Blade and Vane Time Constant 0.5 Platform Time Constant 1 Design Tip Clearance % 1.5 d Flow / d Tip Clear. 2 d Eff / d Tip Clear. 2 d Surge Margin / d Tip Clear. 5

Table 28: GasTurb inputs to the Compressor Interduct tab under Geometry.

Number of Struts 12 Length/Inlet Inner Radius 0.75 Inner Annulus Slope@Exit deg 10 Relative Strut Length % 60 Casing Thickness m 0.005 Casing Material Density kg/m3 2710 Compr Interduct Mass Factor 1

37 Table 29: GasTurb inputs to the HP Compressor tab under Geometry.

Number of Stages 6 Number of Radial Stages 0 Number of Variable Guide Vanes 1 Inlet Guide Vanes (IGV) 0/1 1 IGV Profile Thickness % 5 IGV Material Density 2710 Annulus Shape Descriptor 0…1 1 Given Radius Rat: Inl/Exit 0/1 0 Inlet Radius Ratio 0.5 Exit Radius Ratio 0.9 Blade and Vane Sweep 0/1 0 First Stage Aspect Ratio 2 Last Stage Aspect Ratio 2.6 Blade Gapping: Gap/Chord 0.2 Pitch/Chord Ratio 0.5 Disk Bore / Inner Inlet Radius 0.38 Diffuser Area Ratio 2 Rel Thickness Inner Air Seal 0.04 Compressor Mass Factor 1 Outer Casing Thickness m 0.005 Outer Casing Material Density kg/m3 2710 Casing Thickness m 0.005 Casing Material Density kg/m3 2710 Rel Work of Radial End Stage 0.3 Duct Inner Radius Ratio 1 Duct Length/Inlet Inner Radius 0 Number of Duct Struts 8 Relative Duct Strut Length % 60 Rad Diffuser/Rotor Blade Length 0.5 Rotor Inlet Swirl Angle 0 Rotor Blade Backsweep Angle 20 Diffuser Wall Thickness m 0.0025 Casing Thermal Exp Coeff E-6/K 10 Casing Specific Heat J/(kg*K) 500 Casing Time Constant 10 Blade and Vane Time Constant 0.5 Platform Time Constant 1 Design Tip Clearance % 1.5 d Flow / d Tip Clear. 2 d Eff / d Tip Clear. 2 d Surge Margin / d Tip Clear. 5

Table 30: GasTurb inputs to the Burner tab under Geometry.

Reverse Flow Design (0/1) 0 Outer Casing Length/Length 6 Exit/Inlet Radius 1.15 Length/Inlet Radius 1.1 Can Width/Can Length 0.4 Inner Casing Thickness m 0.005 Outer Casing Thickness m 0.005 Casing Material Density kg/m3 1601.85 Can Wall Thickness m 0.005 Can Material Density kg/m3 1601.85 Can Thermal Exp Coeff E-6/K 10 Can Specific Heat J/(kg*K) 500 Can Constant 1 Mass of Fuel Inj. / Fuel Flow 2 Burner Mass Factor 1

38 Table 31: GasTurb inputs to the HP Turbine tab under Geometry.

Number of Stages = 1 no input Unshrouded/Shrouded Blades 0/1 1 Inner Radius: R,exit / R,inlet 1 Inner Annulus Slope@Inlet deg 15 Inner Annulus Slope@Exit 15 First Stage Aspect Ratio 1 Last Stage Aspect Ratio 1.5 Blade Gapping: Gap/Chord 0.25 Pitch/Chord Ratio 1.88 Disk Bore / Inner Inlet Radius 0.3 Rel Thickness Inner Air Seal 0.04 HP Turbine Mass Factor 1 Outer Casing Thickness m 0.005 Outer Casing Material Density kg/m³ 1922.22 Casing Thickness m 0.005 Casing Cooling Effectiveness 0.5 Casing Material Density kg/m³ 1922.22 Casing Thermal Exp Coeff E-6/K 10 Casing Specific Heat J/(kg*K) 500 Casing Time Constant 20 Blade and Vane Time Constant 2 Platform Time Constant 5 Design Tip Clearance % 1.5 d Eff / d Tip Clear. 2

Table 32: GasTurb inputs to the Turbine Interduct tab under Geometry.

Number of Struts 0 Exit/Inlet Inner Radius 1.45 Length/Inlet Inner Radius 0.45 Inner Annulus Slope@Inlet deg 5 Inner Annulus Slope@Exit deg 35 Relative Strut Length % 40 Casing Thickness m 0.00508 Casing Material Density kg/m3 2710 Turbine Interduct Mass Factor 1

Table 33: GasTurb inputs to the LP Turbine tab under Geometry.

Number of Stages = 7 no input Unshrouded/Shrouded Blades 0/1 1 Inner Radius: R,exit / R,inlet 1.4 Inner Annulus Slope@Inlet deg 33 Inner Annulus Slope@Exit -12 First Stage Aspect Ratio 2.02 Last Stage Aspect Ratio 2.02 Blade Gapping: Gap/Chord 0.95 Pitch/Chord Ratio 0.708 Disk Bore / Inner Inlet Radius 0.5 Rel Thickness Inner Air Seal 0.04 LP Turbine Mass Factor 1 Casing Thickness m 0.00508 Casing Cooling Effectiveness 0.5 Casing Material Density kg/m³ 2710 Casing Thermal Exp Coeff E-6/K 23 Casing Specific Heat J/(kg*K) 900 Casing Time Constant 20 Blade and Vane Time Constant 2 Platform Time Constant 5 Design Tip Clearance % 1.5 d Eff / d Tip Clear. 2

39 Table 34: GasTurb inputs to the Exhaust tab under Geometry.

Number of Struts 12 Strut Chord/Height 0.75 Strut Lean Angle 7 Gap Width/Height 0.15 Cone Angle deg 40 Cone Length/Inlet Radius 4.6 Casing Length/Inlet Radius 0.5 Bypass Radius/Flange Radius 1 Inner Casing Thickness m 0.002 Outer Casing Thickness m 0.005 Casing Material Density kg/m3 2710 Exhaust Duct Mass Factor 1

Table 35: GasTurb inputs to the Bypass tab under Geometry.

Number of Struts 0 Flat Point Pos in % of Length 85 Flat Point Radius/Inlet Radius 1.15 Nozzle Thr Pos in % of Length 90 Strut Inlet Pos in % of Length 60 Relative Strut Length % 10 Mean Strut Thickness m 0.00508 Strut Material Density kg/m3 1601.85 Inner Casing Thickness m 0.002 Outer Casing Thickness m 0.005 Casing Material Density kg/m3 2710 Bypass Duct Mass Factor 1

Table 36: GasTurb inputs to the Nozzle tab under Geometry.

Std/Plug/Power Gen Exh 1/2/3 2 Inl Section Length/Outer Radius Conv Length/Inl Section Radius Cone Angle deg Cone Length/Inlet Radius 2 Inlet Section Area Ratio 0.936 Inner Casing Thickness m 0.00635 Outer Casing Thickness m 0.00635 Casing Material Density kg/m3 1922.22 Nozzle Mass Factor 1

12 Appendix C - Material Properties

Table 37: Properties for Compressor Materials (*Properties similar to nt Al-Fe; **Properties taken at 760 K) [7], [5], [2].

Weldalite 049* Inconel 706B** Titanium Alloy Density (kg/m3) 2710 8050 4430 Ultimate Tensile Strength (MPa) 710 690 900 Elongation % 5 20 10 Yield Tensile Strength (MPa) 690 660 830

40 References

[1] url: en.wikipedia.org/wiki/Ceramic_matrix_composite. [2] url: www.dierk-raabe.com/titanium-alloys/mechanical-properties-of-titanium/. [3] Boeing 787-8 Dreamliner: Operating Manual and Checklists. July 2015. url: wiki . flightgear . org/Boeing_787-8_Dreamliner:_Operating_Manual_and_Checklists. [4] Meherwan P. Boyce. “Chapter 7 - Axial-Flow Compressors”. In: Gas Turbine Engineering Handbook. 4th Edition. Elsevier, 2012, pp. 303–355. isbn: 978-0-12-383842-1. url: www.sciencedirect.com/ science/article/pii/B978012383842100007X. [5] Joachim Kurzke. GasTurb. 2018. url: www.gasturb.de. [6] Qiang Li et al. “High-Strength Nanotwinned Al Alloys with 9R Phase”. In: Advanced Materials 30.11 (2018), p. 1704629. doi: 10.1002/adma.201704629. eprint: onlinelibrary.wiley.com/doi/ pdf/10.1002/adma.201704629. url: onlinelibrary.wiley.com/doi/abs/10.1002/ adma.201704629. [7] Material Property Data. url: www.matweb.com. [8] Pratt and Whitney PW1100G Geared Turbofan Engine. July 2013. url: theflyingengineer.com/ flightdeck/pw1100g-gtf/. [9] Pratt & Whitney GTF Engine. url: www.pw.utc.com/products-and-services/products/ commercial-engines/Pratt-and-Whitney-GTF-Engine/. [10] Researchers develop high-strength nanotwinned aluminum alloy. Jan. 2018. url: www.greencarcongress. com/2018/01/20180126-nt.html. [11] Andrew Rolt et al. “Scale eﬀects on conventional and intercooled turbofan engine performance”. In: The Aeronautical Journal 121.1242 (2017), pp. 1162–1185. url: dspace.lib.cranfield.ac.uk/ bitstream/handle/1826/12185/Scale_effects_on_conventional_and_intercooled_ turbofan_engine_performance-2017.pdf. [12] Titanium and Titanium Alloys: Fundamentals and Applications. Wiley-VCH, 2003.