ME 404: Gas Turbines Team 7 Final Report Nick Rados, Karan Sandhu, Ryan Cranston, Sean Fitzpatrick

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1 ME 404: Gas Turbines Team 7 Final Report Nick Rados, Karan Sandhu, Ryan Cranston, Sean Fitzpatrick NEO 1.0 Acknowledgments Our team used a compilation of references including the textbook, Elements of Propulsion by Jack Mattingly, Gas Turbine Theory sixth edition, specifications of the Rolls Royce Trent 700 and other similar engines from online sources, as well as our calculations from previous home works in order to determine the following specifications of our design. Operating Conditions and Engine Performance Take-off: Mass flow rate: kg/s By-pass ratio (BPR): 7.0 Net Thrust: 70, lbf Specific fuel consumption: kg/hrn T4 = 1500 K Cruise: Altitude: 12,500m Air speed: 575 mph ( m/s) Mach #: 0.83 Mass flow rate: kg/s By-pass ratio (BPR): 7.0 Net Thrust: 17, lbf Specific fuel consumption: kg/hrn T4 = 1500 K Engine Architecture Cruise: Engine overall pressure ratio: TO: Engine overall pressure ratio: Fan-pressure ratio (FPR): 1.70 Compressor pressure ratio (CPR): shaft design: LP- shaft at 3,300.00rpm (Fan/LPT) IP-shaft at 7, rpm (IPC/IPT) HP-shaft at 12, rpm (HPC/HPT) See Figure 7: Engine flow path diagram Stage pressure ratio: Fan: IPC: HPC: LPT: IPT: HPT: # of Stages: Fan Stages: 1 LPC Stages: 0 IPC Stages: 8 HPC Stages: 6 LPT Stages: 5 IPT Stages: 1 HPT Stages: 1

2 Velocity Triangles Compressor Figure 1: Compressor Velocity Triangles Turbine Figure 2: Turbine Velocity Triangles See Compressor_Calc.xlsx and Turbine_Calc.xlsx for calculations respectively.

3 Annulus radii: NEO 1.0 has been specifically designed to fit the AirBus A330 as seen in Figure 3 below. The wing architecture of the A330 has placed size constraints on the engine. Engine parameters are as follows: Fan diameter: 100in (2.54m) Length: 160in (4.064m) Flow area: The tip, hub and mean radii are then calculated using the equations below: Compressor: Hub radius: 0.36 m Tip radius: 0.51 m Mean radius: 0.43 m Hub/tip ratio = 0.70 Turbine: Hub radius: 0.36 m Tip radius: 0.45 m Mean radius: 0.41 m Hub/tip ratio = 0.80 Performance Analysis: See figures in appendix. Mission Fuel burn for 5,000 mile flight is calculated to 16, gallon per mission. Calculations based off airbus A330 as seen in Engine_Calc.xlsx. Loss and Efficiency For our aircraft design, we made certain assumptions for the efficiencies which are shown in the table below. Using Aircraft Engine Design by John D. Mattingly as a reference, we chose an advanced level of technology (level 4) to assume efficiencies which represents typical values for the time period 2005-present. Therefore the efficiencies are at their optimum values since the latest technology is considered. The efficiencies are polytropic efficiencies which are defined as actual as compared to ideal value. In addition, there are pressure losses that arise due to viscous losses in the combustion chamber and total pressure loss due to combustion at finite Mach number. Table 1: Efficiency assumptions

4 Component Efficiency Diffuser Mechanical Compressor 0.9 Fan 0.89 Burner 0.99 Turbine 0.89 Pressure loss 0.05 Nozzle 0.98 We also made some underlying assumptions for the performance analysis: 1. The total pressure ratios do not change from their reference values. 2. The efficiencies listed in the table above do not change from their reference values. 3. Turbine leakage effects are neglected and no power is removed from the turbine to drive accessories. 4. The flow on average is steady and one-dimensional at the entry and exit of each component. 5. Effect of cooling will reduce turbine efficiency from approximately 0.91 to The cooling flow mixes perfectly with the flow in the turbine. 7. Fluids will be assumed to behave as a perfect gas, calorically perfect, with constant molecular weight across the diffuser, fan, compressor, turbine, nozzle, and connecting ducts. Cooling System With modern gas turbine engines operating between 1473K and 1873K, turbine blade cooling is an absolute necessity. The metals that are used on modern-day turbine blades are not capable of withstanding the high temperatures that are experienced; therefore, they would ultimately melt if not properly cooled. Cooling methods can be divided into convection cooling, impingement cooling, film cooling, full-coverage film cooling, and transpiration cooling. [M., Lunai, ] Our combustion temperature of 1500K is over the limit and therefore cooling is required. For our engine design, we plan to incorporate both internal and external cooling through impingement cooling and film cooling, respectively, in our turbine blades in order to ensure that there will be no failure of the turbine blades themselves. Impingement cooling works by injecting a coolant in the form of impingement jets within the wall of the turbine that will increase the heat exchanger capabilities. As displayed in Figure 3 below, the turbine blades are hollow with a core and internal cooling passages. The air enters from the front end of the blade and flows towards the

5 back. We know that impingement cooling is a far more effective system than film cooling, because it allows more heat to be transferred; however, we plan to additionally incorporate film cooling in order to maximize our cooling performance. Film cooling works by using the cold air that is bled from the compressor stages, injecting the air into the blade, resulting in the air being pushed out through small holes in the blade walls. The air that is discharged from the holes will provide a protective film along each turbine blade. Many factors are taken into account when determining the performance, such as the injection geometry, surface geometry and roughness, the amount of turbulence, the coolant properties, and lastly the amount of mass flow. A visual representation of how the film cooling works is displayed in Figure 4 below. Figure 3: Impingement Cooling Figure 4: Film Cooling Addition of cooling flow will increase the mass flow rate and this will be accommodated by an increase in hub radius. 5% of the core flow is the cooling air at 1500K. Applying the conservation of mass equation, we calculated the change in annulus dimensions with the addition of cooling due to increased mass flow rate. There was a small change in the annulus radius due to increased mass flow rate resulting in an increased annulus area which in turn increased the annulus radius. These changes can be seen in Figures 5&6 below. (Coolingflowanalysis.xlsx)

6 Figure 5: Cooling flow analysis Figure 6: Close up of analysis showing radius change Engine Construction: Estimated weight: 10,600lbs projected from Trent 700. Compressor blades are primarily constructed from Alloy 834 Titanium. [Bayer, ] Traditionally combustor components are fabricated out of sheet nickel-base super-alloys. Our turbine hardware will be constructed with these alloys along with HA-188, a cobalt base superalloy. Shafts will be constructed of 1%Cr-1.25%Mo-0.25%V Steel in hardened and tempered condition. Polymer matrix composites will be utilized for the exterior of the engine to minimize weight while maintaining strength. [Schilke, 2004]

7 Appendix Figure 7: Engine flow path diagram Figure 8: Pressure through Engine Figure 9: Temperature through Engine

8 SFC [kg/(h*n)] FPR = 1.5 FPR = 1.6 FPR = 1.7 FPR = BPR Figure 10: FPR vs. BPR vs. Cruise SFC Figure 11: OPR vs. T4 vs. Cruise SFC

9 Figure 12: OPR vs. FPR vs. Cruise SFC Figure 13: BPR vs. T4 vs. Cruise SFC

10 References 1. Mattingly, Jack D., and Hans Von Ohain. Elements of Propulsion: Gas Turbines and Rockets. Reston (Virginia): American Institute of Aeronautics and Astronautics, Print. 2. Mattingly, Jack D., William H. Heiser, and Daniel H. Daley. Aircraft Engine Design. 2nd ed. Washington, D.C.: American Institute of Aeronautics and Astronautics, Print. 3. M., Luai, S.m. Shaahid, and Ali A. "Jet Impingement Cooling in Gas Turbines for Improving Thermal Efficiency and Power Density." Advances in Gas Turbine Technology (2011): Web. 4. "Rolls-Royce Trent 900 Engines Provide Power for First A380." -- LONDON, January 16 /PR Newswire UK/ --. Rolls-Royce, n.d. Web. 01 Dec "Turbine Blade Cooling." Turbine Blade Cooling. EPFL, 19 Dec Web. 01 Dec < 6. "What Is Film Cooling?" What Is Film Cooling? N.p., n.d. Web. 01 Dec "Why Hasn't Commercial Air Travel Gotten Faster since the 1960's?" MIT: School of Engineering. MIT, 19 Feb Web. 02 Dec "Airbus A330 Medium to Long-Range Jetliner." AerospaceWeb. N.p., n.d. Web. 6 Dec Bayer, R.R. An Overview on the Use of Titanium in the Aerospace Industry, Materials Science and Engineering A, Vol.A213, (1996), pp ISSN Schilke, P.W. (2004) Advanced Gas Turbine Materials and Coatings,