IMPROVING ENGINE EFFICIENCY THROUGH CORE DEVELOPMENTS

Brief summary: The NASA Environmentally Responsible Aviation (ERA) Project and Fundamental Aeronautics Projects are supporting compressor and turbine research with the goal of reducing aircraft engine fuel burn and greenhouse gas emissions. The primary goals of this work are to increase aircraft propulsion system fuel efficiency for a given mission by increasing the overall pressure ratio (OPR) of the engine while maintaining or improving aerodynamic efficiency of these components. An additional area of work involves reducing the amount of cooling air required to cool the turbine blades while increasing the turbine inlet temperature. This is complicated by the fact that the cooling air is becoming hotter due to the increases in OPR. Various methods are being investigated to achieve these goals, ranging from improved compressor three-dimensional blade designs to improved turbine cooling hole shapes and methods. Finally, a complementary effort in improving the accuracy, range, and speed of computational fluid mechanics (CFD) methods is proceeding to better capture the physical mechanisms underlying all these problems, for the purpose of improving understanding and future designs.

National Aeronautics and Space Administration

Improving Engine Efficiency Through Core Developments

Dr. James Heidmann Project Engineer for Propulsion Technology (acting) Environmentally Responsible Aviation Integrated Systems Research Program

AIAA Aero Sciences Meeting January 6, 2011 www.nasa.gov NASA’s Subsonic Transport System Level Metrics …. technology for dramatically improving noise, emissions, & performance

N+1 = 2015 N+2 = 2020 N+3 = 2025 CORNERS OF THE Technology Benefits Relative Technology Benefits Relative Technology Benefits TRADE SPACE To a Single Aisle Reference To a Large Twin Aisle Configuration Reference Configuration

Noise -32 dB -42 dB (cum below Stage 4) -71 dB

LTO NOx Emissions -60% -75% better than -75% (below CAEP 6)

Performance: -33% -50% better than -70% Aircraft Fuel Burn Performance: -33% -50% exploit metro-plex* concepts Field Length

Goals are relative to reaching TRL 6 by the timeframe indicated Engine core research primarily focused on fuel burn metric (SFC) Core developments have positive and negative impacts on NOx

2 POTENTIAL REDUCTION IN FUEL CONSUMPTION Advanced N+2 Configurations

Advanced Configuration #1 Advanced Configuration #2B Advanced Configuration #2A N+2 HWB300 N+2 “tube-and-wing“ N+2 HWB300 2025 EIS (TRL=6 in 2020) 2025 EIS (TRL=6 in 2020 assuming 2025 EIS (TRL=6 in 2020) accelerated technology development)

Advanced Engines -120,300 Δ Fuel Burn = -139,400 lbs lbs -15.3% (-49.8%) (-43.0%) -151,300 lbs Advanced Advanced (-54.1%) Engines Engines Δ Fuel Burn = Δ Fuel Burn = -16.0% -18.5%

Embedded Engines with Fuel Burn = 159,500 lbs BLI Inlets ∆ Fuel Burn = -3.2% -43.0% Fuel Burn = 140,400 lbs Fuel Burn = 128,500 lbs -49.8% -54.1% Propulsion Technology Enablers

Fuel Burn - reduced SFC (increased BPR, OPR & turbine inlet temperature, potential embedding benefit)

W Aircraft Velocity Lift fuel = ln 1 + Range TSFC Drag WPL + WO

• Engine Fuel •Aerodynamics • Empty Weight Consumption

TSFC = Velocity / (ηoverall)(fuel energy per unit mass)

ηoverall = (ηthermal)(ηpropulsive)(ηtransfer)(ηcombustion)

assuming constant component efficiencies

Core research impacts thermal efficiency through increased OPR High power density cores enable higher propulsive efficiency cycles Low pressure turbine improvements impact transfer efficiency Propulsion Technology Opportunity

Propulsion system improvements require advances in both propulsor and core technologies

Core Improvements (direct impact on LTO NOx)

Propulsor improvements Cycle Performance Improves with Temperature

6 Engine Thermal Trends

1800

HYPRB 1700 M-88 AERO-ENGINE XG40 HYPRA 1600 F100-PW-299

] INDUSTRIAL GT Temp. at Rotor Inlet GE90 701G ℃ PW4000 501G 1500 F110 V2500 TRENT F404 9001G,H F101 7001G,H E3 AGTJ100B 1400 CFM56-5 F100 PW2037 CF6-50 AGTJ100A 701F CF6-6 CFM56 CF6-80 501F CGT 1300 TF39 RB199 9001F TEPCO TF30 JT9D-7R4 H-25 7001F RB211-22 FJR710/600 MF111 1200 FJR710/20 JT9D-3 701D T56-14 TF40 501D 7001E 1100 JT8D FJR710-10 9001E JT15D 9001B Conway Avon J79-17 501B XF3-20 F3-30 JT3D 1000 7001B TFE731 JR220 J69 Dart10 T64 900 CT610 J3-7 501AA J73 BMW003 JR100 800 W2/700 J3 JUMO004-1 700 HeS3B W1 NE-20

TURBINEINLET TEMPERATURE [ F.Whittle 1GO 600 PAT BBC4000kW BBC900kW 500 KO-7 1930 1940 1950 1960 1970 1980 1990 2000 YEAR

From Dr. Toyoaki Yoshida, National Aerospace Laboratory, Japan Turbine Materials Improvements

Increase in operational temperature of turbine components. After Schulz et al, Aero. Sci. Techn.7:2003, p73-80. 8 Turbine Cooling Improvements

9 Turbomachinery Aero Design-Based Tech Enablers

Low-Shock Aspiration Flow Design, High Controlled, Efficiency, Highly-Loaded, High Pressure Low Pressure Turbine Turbine

Highly-Loaded, Low Pressure Multistage Turbine Compressor (higher Plasma Flow efficiency and OPR) Control

FLOW

UW

FLOW

Novel Turbine z Cooling Concepts

High-Efficiency Centrifugal Compressor (small high efficiency core)

10 Multi-Stage Axial Compressor (W7)

Objective: To produce benchmark quality validation test data on a state-of-the-art multi-stage axial compressor featuring swept axial rotors and stators. The test in ERB cell W7 will provide improved understanding of issues relative to optimal matching of highly loaded compressor blade rows to achieve high efficiency and surge margin.

ERB Test Cell W7 Approach: Test a modern high OPR axial compressor representative of the front stages of a commercial engine high pressure compressor in partnership with General Electric. Test will enable improved high OPR designs for reduced engine SFC. NASA 3-Stage Axial Compressor

11 UTRC NRA – High Efficiency Centrifugal Compressor (HECC)

m = 10.1 lbm/s 90.0% 89.5% Engine 89.0% scale Opportunity for improved rotary wing 88.5% polytropic vehicle engine 88.0% efficiency is estimated performance as well 87.5% as rear stages for as 87.9 - high OPR fixed wing efficiency polytropic TT 87.0% 88.9% application 86.5%

86.0% CFD prediction apply delta, CFD increase inlet scale from rig to to data radius ratio engine 12 Turbine Film Cooling Experiments

Objective: Fundamental study of heat transfer and flow field of film cooled turbine components Rationale: Investigate surface and flow interactions between film cooling and core flow for various large scale turbine vane models Approach: Obtain detailed flow field and heat transfer data and compare with CFD simulations

Large Scale Film Hole: Film cooling jet downstream of hole

Trailing Edge Film Vane Heat Transfer: Ejection: Good agreement between IR images GlennHT and experiment Anti-Vortex Film Cooling Concept

Flow Direction

Front View

Top View

Side View

Auxiliary holes (yellow) produce counter- vorticity to promote jet attachment Comparison of round hole and “anti-vortex” Advantages: Inexpensive due to use of only turbine film cooling jet attachment round holes, hole inlet area unchanged 14 NASA/General Electric Highly-Loaded Turbine Tests

4 TON HOIST

COFFING

FLOW

FLOW

Turbine Testing in NASA Glenn Single Spool Turbine Facility (W6)

Unique High-Speed High Pressure Ratio Capability

15 NASA/General Electric Highly-Loaded Turbine Tests

Conventional HPT Reduced Shock Design

Pressure Ratio (PTR/PS) = 3.25 HPT: Reduced Shock Design Stage Pressure Ratio = 5.5 LPT: Flow-Controlled Stator & Contoured Endwall

Enables efficient high overall pressure ratio turbine capability with reduced cooling flow and reduced SFC 16 Dielectric Barrier Discharge Plasma Actuators

Low pressure turbine flow control – reduced weight and improved efficiency

DBD PLASMA Advantages of GDP actuators: ELECTRODE “WIND” • Pure solid state device

APPLIED • Simple, no moving parts INSULATOR VOLTAGE • Flexible operation, good for varying operating conditions ELECTRODE • Low power • Heat resistance – w/ proper materials

F U Force Versus Pulse Repetition Rate & Bias L Electrode perpendicular Force, mN/m O W 12 to flow 11 W Bias Voltage, 40 10 p-to-p, kV 9 Active Flow Control via 0 8 1.4 1.500 30 7 2.7 3.000

4.0 4.500 6 Oscillating wall jet 5.3 6.000

5 6.7 7.500

8.0 PRR,kHz

4 20 9.000 Force, mN/m Force, 9.3 10.50 F 3 10.7 12.00 L z 2 11.0 Electrode parallel to flow 1 10 O 0 0 10 20 30 40 50 2.7 5.3 8.0 10.7 W Active Flow Control via PRR, kHz Princeton Nanosecond PulsingBias Voltage, NRA p-to-p, kV Streamwise vortices Large force induced with voltage bias 17 CMC Engine Components Reduce Cooling Air Requirements

High Pressure Low Pressure Combustor Exhaust Nozzle Turbine Turbine Temperature 2200-2700°F 2400-2700°F 2200-2300°F 1500-1800°F CMC System SiC / SiC SiC / SiC SiC / SiC Oxide / Oxide • Reduced cooling • Reduced cooling • Reduced cooling • Light weight Engine Benefit • Reduced NOx • Reduced SFC • Strength / weight • Noise reduction • Pattern Factor • Higher use temp • Durability • Manufacturing • Manufacturing • Manufacturing • Attachment & • Durability • Durability Challenges Integration • Attachment & • Attachment & • Durability Integration Integration

18

1 8 CMC Turbine Vane Reduces Fuel Burn

Prepreg lay-up assembly • Hi-Nic type S fibers • BN interface coatings • Balanced ply lay-up • 0/90o tapes • Fiber volume ~ 28%

CVI SiC with MI SiC • Hi-Nic Type S fibers • CVI BN fiber coatings • 5 harness satin weave • Fiber volume ~ 35%

Durability comparison of candidate CMC material systems planned for 2011

19 CMC Nozzle Reduces Weight, Increases Temperature Capability, Potential Noise Benefit • NASA teaming with Rolls Royce/LibertyWorks on CMC exhaust mixer nozzle development • Subscale aero-rig component testing (<12” dia.) • Example of a similar article fabricated by ATK COIC shown. • Structural benchmark testing at NASA GRC, with stress & failure model validation to follow.

18-inch dia. CMC Mixer Demonstration Article

CMC (Ceramic Matrix Composite) • ATK COIC Oxide/Oxide CMC: AS-N610 (Aluminosilicate matrix, Nextel 610 fabric reinforcement) • Composition: 51% fiber, 24% matrix, 25% open porosity

20 Core Engine Research Summary

Core turbomachinery research directly impacts fuel burn reduction goals of ERA and other NASA Aeronautics projects

Compressor research focused on increasing overall pressure ratio while maintaining or improving aerodynamic efficiency

Turbine research focused on increased loading, reduced cooling flows, and improved aerodynamic efficiency

High OPR axial compressor testing with General Electric

Centrifugal compressor testing with United Technologies Research Center

Highly-loaded HPT testing with General Electric

Fundamental testing of turbine cooling flows and low pressure turbine flow control with universities and Department of Energy

Computational fluid dynamic development and assessment across all components, including advanced turbulence models such as LES and DNS