Advanced Propulsion System Design: Applications of CFD J. J. Isaac Head, Propulsion Division National Aerospace Laboratories Foundation Day Lecture: C- MMACS 15 TH September, 2006
Propulsion Trends Scramjet Ramjet Afterburning Turbojet Low Bypass Turbofan with Afterburner Low Bypass Turbofan High Bypass Turbofan Prop-fan Turboprop Piston-prop Propulsion innovations have been the fundamental drivers to progress in air transportation 0 1 2 3 4 5 6 Mach Number
Propulsion Division R&D Activities Hypersonic Air Breathing Propulsion (Ramjet & Scramjet Combustor Technology) Advanced Gas Turbine Technology
Anderson s three dimensions of modern fluid dynamics Pure Experiment Pure Theory Computational Fluid Dynamics (CFD) Prandtl was able to see the solutions of the differential equations without calculating them ---- Werner Heisenberg Prandtl s rule Clean Aerodynamics Everyone believes experimental results except the person who performed the experiment and no one believes the numerical results except the person who performed the prediction!
C F D Advantages CFD can be used as a design tool CFD can be used as a research tool Provide virtual solutions for easy review / investigation Provide valuable flow visualization to explain experimental observations Excellent complementary tool to testing Issues: Prediction of transition from laminar to turbulent flow Calculation of separated flow and unsteady wakes Turbulence modeling Chemical Kinetics Grid Generation
Turbomachinery: Propulsion Technology Challenges - Fan, compressor, high-pressure turbine, low pressure turbine efficiency improvements through CFD and experimentation -Inlet, engine, control system compatibility; operability - Turbine life and cooling - Component reaibility Combustor - Development of a viable commercial design; new fuel injection, mixing, variable geometry, and high temperature walls with new ceramic matrix composite (CMC) materials -Validation of technology readiness - Combustor reliability and life Jet nozzle performance, acoustic, durability Materials Advanced digital controls Weight Manufacturing technology (Ref: Aeronautical Technologies for the 21 st Century: Committee on Aeronautical Technologies, Aeronautics and Space Engineering Board, National Research Council, USA)
High-Speed Computation for Propulsion! CFD and high-speed computation in general will be essential for new engine development in the coming decades, in order to shorten the time for development and to discover new design optima! For the foreseeable future in Propulsion, both CFD and experiments will be needed to efficiently apply results to meet design goals! Analysis codes are necessary to provide the basic predictive ability, at the finest level of detail, for the physics and chemistry underlying propulsion! Design codes are the essential direct tools for engine development; they should embody the best capabilities of analysis codes but, because of their different purpose, be far faster, easier to apply and therefore, necessarily less comprehensive and precise! Technical challenges for high-speed computation include " multistage turbomachinery flows with mixing and shock losses; " combustor design for low emissions, using codes that integrate chemical kinetics, fluid mechanics, heat transfer, materials and structures; " unsteady phenomenon fully represented in codes pertaining to new ideas about the stability and control of propulsion systems; and " acoustics and noise, which must- through CFD - be featured in the primary design process (Ref: Aeronautical Technologies for the 21 st Century: Committee on Aeronautical Technologies, Aeronautics and Space Engineering Board, National Research Council, USA)
Kaveri Engine for the LCA
Ignition and Combustion in an Afterburner and Main Combustor The problem of flame stabilization Flow velocities >> Flame propagation velocities of conventional fuels (kerosene, hydrogen) In a scramjet through flow velocities are over a 1 km/s, akin to lighting a match inside a tornado and somehow keeping it alight Fuel / air ratio Sheltered zone Heat loss
Hardware Quasi-morphological study CFD Study -Temperature contour at X/L = 0.36 plane (CFD ACE +) Kaveri ring-radial afterburner
Kaveri Full-Scale Afterburner Tests Full Scale AB Test Rig Flame Stabilizer & fuel ring assembly AB end-on view AB side view
CFD ANALYSIS OF GAS TURBINE COMBUSTOR : A CAD TO MESH TO CFD APPROACH Complete Sector View of Combustor
Animation of Velocity Vectors
Animation of Temperature Contours Axial Direction
Total Pressure Loss ( % ) PHOENICS Expt. CFD-ACE Pre Diff. Exit 0.15-0.07 Outer Annulus Dump Diff. 1.65-1.79 Inner Annulus Dump Diff. 1.33-1.77 Exit 5.61 4.24 4.42 Mass Flow Analysis ( % ) PHOENICS EXPT. Cold Flow (CFD-ACE) Hot Flow (CFD-ACE) Outer 46.10 43.24 42.77 43.28 Inner Annulus 41.12 40.86 40.39 40.73 Core 12.78 15.90 16.84 15.99
CFD as a Combustor Design Tool
K4 COMBUSTOR K9 MODIFIED COMBUSTOR
Enlarged Velocity Vectors near Primary Zone K4 Combustor K9 Modified
AVAILABLE CODES FOR TURBOMACHINERY FLOW ( 3D VISCOUS FLOW IN TURBOMACHINERY) DAWES CODE: un_b3d_ke Unsteady/Steady Flow, Structured H-Grid, Multistage, Options for Mixing Plane Treatment GENERAL PURPOSE CFD CODE: CFD-ACE+ - 3-D Navier-Stokes general purpose CFD code
3-Stage LP Compressor of a Gas-Turbine Engine Grid: 33(radial)x331(axial)x33(tangential) Fan axis Fig 2.0.1b Grid for 3-stage compressor Design Pressure Ratio=3.4
M At mid-pitch At mid-height Case A2 Fig 4.1.2. Mach number contours on some planes.
3.5 LPC perfermance map Pressure ratio: 3.4 3.4 3.3 3.2 Design speed 95% speed 90% speed A2 Design speed: 10312 rpm Pr. ratio 3.1 3 2.9 2.8 2.7 2.6 2.5 60 62 64 66 68 70 72 74 76 78 mass flow in kg/s Pressure ratio Vs. Mass flow Efficiency Vs. Mass flow 90 LPC performance map 88 86 84 efficiency 82 80 78 76 Design speed 95% speed 90% speed Performance map for 3-stage compressor configuration 74 72 70 60 62 64 66 68 70 72 74 76 78 mass flow in kg/s
AFCR(NAL) COMPRESSOR Rotor Stator AFCR(NAL) COMPRESSOR Pressure ratio 1.35 Design speed 12930 rpm Nominal tip clearance 0.8mm Number of blades (R/S) 21/18 Grid size # 33 x 246 x 33 Without Tip clearance ( 2,67,894 Cells) # 33 x 246 x 37 With Tip clearance (3,00,366 Cells) #9 Grids within the tip clearance Transonic AFCR compressor stage and grid
With 0.0mm TC With 0.8mm TC Velocity vectors on blade to blade plane near casing Comparison of Experimental and Predicted Performance for Transonic Axial Flow Compressor
CDNAL BLADE CASCADE CALCULATIONS Blade-to-blade Hub-to-tip Grid 31x91x31 Surface Mach number, M 1.1 1.0 0.9 0.8 0.7 0.6 0.5 0.4 P-A(1.21)-I(43.7) S-A(1.21)-I(43.7) P-A(1.26)-I(43.7) S-A(1.26)-I(43.7) P-A(1.30)-I(43.7) S-A(1.30)-I(43.7) TCT-P-A(1.26)-I(43.7) TCT-S-A(1.26)-I(43.7) 0.3 0.2 0.1 0.0 0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 X/Cax Fig. Comparison of surface Mach number for CDNAL blade (AVDR=1.21,1.26,1.30; Grid=31x91x31)
DESIGN OF CDA PROFILE SECTION-1 45 40 35 30 Section 1 MCA CDA d3 MCA and CDA-3 profiles 25 Y 20 15 10 CDA Design-3 5 0-5 MCAD3S1.GRF -5 0 5 10 15 20 25 30 35 40 45 X 1.2 sec1 1 CDA-3 MCA 0.8 Surface Mach number distribution Mach no 0.6 0.4 0.2 0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c
LP Compressor performance* with CD design for stator 3 90 Kaveri LPC efficiency 88 86 84 MCA CDA-2 CDA-3 Efficiency Vs Mass flow 82 80 78 65 67 69 71 73 75 77 79 81 Mass flow in kg/sec 4 Kaveri LPC 3.9 3.8 Pressure ratio 3.7 3.6 3.5 3.4 3.3 3.2 MCA CDA-2 CDA-3 Pressure ratio Vs Mass flow 3.1 3 65 67 69 71 73 75 77 79 81 Mass flow in kg/sec * GTRE Results
1.4 GTRE KAVERI ENGINE FAN - MCA PROFILE OF IIIrd STATOR Surface Mach Number Distribution M1 = 0.76, M2 = 0.591 & Beta1 = 43 Deg 1.2 1.0 Mach Number 0.8 0.6 0.4 0.2 CFD - Dawes Code TCT - EXPT 0.0 0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 X / Cax
NAL TRANSONIC CASCADE TUNNEL OIL FLOW VISUALISATION MCA Profile Separation Bubble Inlet
LOX First Rotor Cascade Assembly In TCT LOX First Rotor Cascade Assembly
1.8 LPSC - LOX FIRST ROTOR Surface Mach Number Distribution M1 = 0.708 & M2 = 1.097 & Beta1 = 68 Deg 1.6 1.4 1.2 Mach Number 1 0.8 0.6 0.4 0.2 0 CFD - Dawes Code Points - EXPT 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 X / Cax
CODE Three dimensional viscous flow code CFD- ACE+ Pressure-based method SIMPLEC (Semi-Implicit Method for Pressure Linked Equations Consistent). Turbulence model k-ε Mass imbalance as low as 0.0001%. Total - to-total Adiabatic Efficiency(%) Stagnation Pressure Ratio (P 02 /P 01 ) 100 95 90 85 80 75 Without Tip Clearance (0mm) With 3% Tip Clearance (0.291mm) With 4.8% Tip Clearance (0.500mm) With 7.0% Tip Clearance (0.721mm) Where, D = Design point *Constant Tip Clearance (TYPE - I) 70 0 0.5 1 1.5 2 2.5 3 6.5 6 5.5 5 4.5 4 3.5 3 2.5 Where, C = Choke point D = Design point S = Surge point 2 Without Tip Clearance (0mm) 1.5 With 3% Tip Clearance (0.291mm) With 4.8% Tip Clearance (0.500mm) 1 With 7.0% Tip Clearance (0.721mm) Surge Line for 4.8% Designed Tip Clearance 0.5 *Tip clearance is indicated as percent of blade height at exit. *Constant Tip Clearance (TYPE - I) 0 0 0.5 1 1.5 2 2.5 3 S D D m θ0 ( δ )( kg s) Corrected Mass Flow Rate / C m θ 0 ( δ )( kg s) Corrected Mass Flow Rate / Tip Clearance or Gap TYPE I Constant Tip Gap Drop in Peak Efficiency ( η ) (%) 5 4.5 4 3.5 3 2.5 2 1.5 1 0.5 TYPE - I TYPE - II Tip Clearance or Gap ( ) ( ) TYPE II Variable tip gap η= η without tip clearance η with tip clearance 0 0 1 2 3 4 5 6 7 Tip Clearance (% of Blade Height at Exit) Constant tip clearance provides better performance compared to variable tip clearance
Supersonic Combustion ABLV Solid Model of Dual-mode Ramjet/Scramjet Combustor Intake Combustor Isolator Nozzle M=2 Strut base Kerosene / Hydrogen Flame Anchoring in cavity flame stabilizer Schlieren of cavity flame stabilizer Dual Mode Ramjet / Scramjet Combustor & Test rig Kerosene / hydrogen supersonic flame
Computation of supersonic cold cavity flow for scramjet velocity profile at different sections (xc measured from end of step; U/U1 scale for xc=10mm profile) y (m) 0.1 0.08 0.06 0.04 0.02 0-5 0 5 10 15 U/U1 xc=10mm xc=20 xc=30 xc=40 xc=50 xc=60 xc=70 xc=80 xc=90 xc=100 xc=110 xc=120 xc=130 xc=140
Velocity Plot at different sections of slant wall distance from slant wall (m) 0.1 0.08 0.06 0.04 0.02 0-2 0 2 4 6 V/Vinf x=85mm x=88mm x=91mm x=94mm x=97mm Static Pressure Profile at different sections of slant wall Distance from slant wall (m) 0.1 0.08 0.06 0.04 0.02 0 0 5 10 15 P/Pin_air X=85mm X=88mm X=91mm X=94mm X=97mm Computation of supersonic combustion slant cavity flow for Scramjet Exit pressure profile Exit temperature profile
Small Gas Turbine Tomahawk Cruise Missile Distributed Vectored Propulsion
Hybrid Bearing Alternator Compressor FUEL Starting Motor SPECIFICATIONS: Power output-10kw Fuel-Natural Gas/LPG Rotational speed-120000rpm Mass flow rate-0.13kg/s Compressor pressure ratio-3 Turbine inlet temperature-900c Compressor tip Dia-85mm Turbine tip Dia-110 Recuperator Effectiveness-85% Thermal efficiency-30% MGT Turbine Recuperator Gas Generator
Reverse Flow Annular Combustor for 10kW Micro Gas Turbine Application: Stationary Power Unit Specifications: Power output:10kw Fuel: Kerosene/LPG Mass flow rate: 0.13kg/s Compressor pressure ratio: 3 Turbine inlet temperature: 900 C Challenges: Getting satisfactory recirculation patterns Containing the flame inside the flame tube INLET EXIT CAD Model - Cut section Fuel injection Temperature contours Temperature contour just (45 degree segment downstream of atomisers Micro Gas Turbine Combustor
MICRO GAS TURBINE
SMALL GAS TURBINE ENGINE (Straight Jet) (Thrust = 2.25 kn, SFC =1.1 kg/hr/kg) 340 mm diameter* 670 mm length
Cavity Combustor Discrete fuel injection Temperature contours Continuous fuel injection Proposed New Concept: Annular Atomiser Velocity vectors near cavity CAD Model 2D - Cavity Combustor Flame from exit section * Circumferentially uniform Temp. distribution * Compact Combustor * Reduced hot spots
Aerodynamic Cavity Velocity contour animation Velocity vector animation
Cavity Combustor Experiment - Video
Annular Cavity Combustor Proposed Small Gas Turbine Engine CAD Model Test rig assembly CAD Model Annular Cavity Combustor Fabricated part cavity ring Applications: Small Gas Turbine Combustor Un-manned Air Vehicles Cruise Missiles
Concluding Remarks An overview of the state of practice of current CFD tools used for advanced propulsion system design and analysis at NAL has been presented. This included the validation requirements of CFD codes applicable to such advanced propulsion systems. The computational facilities made available by C-MMACS have been invaluable for these studies. Acknowledgements: C-MMACS; special mention Dr. Gangan Pratap, Mr. R. P. Thangavelu, This presentation has been based on the work carried out by the Scientists of the Propulsion Division; special mention Dr. T. R. Shebharkar, Mr. H. S. Muralidhara, Mr. Ashfaque Khan, Mr. Pratap Nayaka, Mr. BBC Kumar, Dr. S. Ramamurthy, Mr. KMM Swamy, Mr. Senthil Kumar, Mr. Dilip Kumar Alone, Mrs. M. T. Shobhavathy, Mr. P. Manjunath, Mr. C. Rajashekar, Mr. Jagannatha Rao, Mr. S. Venkatesh, Mr. V.S. Krishnakumar, Mr. Sathiyamurthy, Mr. Venu, Mr. G.S. Sreenath, Mr. M. Baskaran, Mr. J. Srinivas, Mr. M. Satish Kumar and Mr. ATLN Murthy