Resonant Pulse Combustors: A Reliable Route to Practical Pressure - PowerPoint PPT Presentation

https://ntrs.nasa.gov/search.jsp?R=20170007284 2017-08-12T04:59:11+00:00Z National Aeronautics and Space Administration Resonant Pulse Combustors: A Reliable Route to Practical Pressure Gain Combustion Dan Paxson NASA John H. Glenn Research Center Cleveland, OH International Constant Volume Detonation Combustion Workshop Poitiers, France June 13-16, 2017 www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Acknowledgements This effort summarized in this presentation contains contributions from (and would not have been possible without) the following individuals • Shaye Yungster - CFD • Doug Perkins - Analysis • Scott Jones - Analysis • Kevin Dougherty - Experiments • Robert Pelaez - Experiments • Paul Litke - Experiments • Andy Naples - Experiments • Mark Wernet - PIV • Trevor John - PIV www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Outline • Motivation • Experimental Investigations • Numerical Investigations • Ongoing and Future Directions • Concluding Remarks Pressure Gain Combustion (PGC) Defined: A fundamentally unsteady process whereby gas expansion by heat release is constrained, causing a rise in stagnation pressure and allowing work extraction by expansion to the initial pressure. Context: Our Focus Is Not the Promotion of Any One PGC Mode It Is the Practical Utilization of Confinement www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Motivation  P>0.0, P4/P3>1 Fan Pressure Gain Combustion Theoretically: + Increases thermodynamic cycle efficiency PGC + Reduces SFC / fuel burn ( NASA Objective ) + Reduces greenhouse gas emissions ( NASA Objective ) + Competes with conventional cycle improvements Compressor Turbine Constant Specific Thrust Engine Parameter Turbofan Turbojet 12.0 OPR 30.00 8.00 Turbojet η c 0.90 0.90 10.0 SFC Reduction, %  =1.35 η t 0.90 0.90 Turbofan 8.0 Mach Number 0.80 0.80 T amb (R) 410 410 6.0 T combustor exit (R) 2968 2400 Burner Pressure Ratio 0.95 0.95 4.0 Equivalent to: T sp (lb f -s/lb m ) 18.26 75.86 -6.0% increase in  c SFC (lb m /hr/lb f ) 0.585 1.109 -2.5% increase in  t 2.0 -1 compression stage 0.0 0.95 1.05 1.15 1.25 Low NOX Constraint Combustor Total Pressure Ratio on All Concepts www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Motivation R esonant P ulse C ombustor-RPC (aka ‘Confined’ Volume Deflagration) FEATURES: •Self-sustained operation • No spark plugs •Only one moving part •Relatively low unsteadiness amplitudes • Lower thermal and mechanical stresses • Effluent easier to smooth • Fewer potential issues for downstream turbomachinery •Readily operates with liquid fuels (gasoline, ethylene, kerosene) •Effective lean operation (low T t4 ’s) with bypass ejectors •Unequivocally a pressure gain device • Only known PGC system to operate under static conditions DRAWBACK •Only Modest Pressure Gain is Possible • Confined (not constant) volume combustion Practically: Features May Outweigh Drawback – Even Compared to Other PGC Approaches www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Motivation Spark plug Resonant Pulse Combustion Basic Cycle Starting air Valve Fuel Valve fully closed Valve closing start Valve fully open Valve opening start Combustion Chamber Pressure t x www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Experimental Investigations Ejector Mixing and Pumping Optimization PIV Measured Flowfield Load Cell Ejector Pulsejet Fuel line Thrust plate • 18:1 and greater entrainment ratios • Thrust augmentation ratios up to 2.0 • Velocity fluctuations reduced by 83% Closed Loop Operation in a Gas Turbine Pressure Gain in a Shrouded Configuration Burst Disc Total Pressure gap 1 Fuel Tank Pressurization Line  P Turbocharger Heating Coil Starting Air Line Total Temperature Shroud Total Pressure Total Start Air Temperature Laser Air flow 4 Ejector Pulsejet Optical speed sensor 0 Perforated Liner Load Cell Start Air Burst Disc Oil gap Fuel Tank Pressurization Line Fuel  P Struts Starting Air Line Total Temperature Shroud Total Pressure Static Pressure Thrust plate Static Pressure Air flow Pulsejet Ejector • PR=1.037 @ TR=2.2 Total Pressure Perforated Liner Total Temperature Fuel • rms p ′ /P=4.5% in the shroud Struts 3 Total Temperature 2 Static Pressure • Successful operation at 2 Atm. inlet pressure Total Temperature Static Pressure Total Pressure Static Pressure Total Temperature All Work Done With COTS Hobby Scale Pulse Combustor (Pulsejet) www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Experimental Investigations Spark Off Aux. Air Off Results: 12 120 Thrust, lb f or Fuel Rate, gph •True closed loop operation @ SLS thrust 10 100 fuel • All air supplied by compressor Speed, krpm 8 80 speed •(P Tin /P cout - 1)=3.5% @ T Tin /T Cout =2.2 6 60 •Sustained operation on liquid fuel 4 40 Start • Limited only by COTS reed valve test 2 20 period •Successfully produced thrust 0 0 •Demonstrated Benefit 0 10 20 30 40 1400 Time, sec. T tCin 1300 TCin • Turbine slows and stops with TCout T tCout 1200 Temperature, R conventional combustor at same T Tin /T Cout T tCCin TCCin 1100 TTin T tTin •-20 dB noise reduction across Turbine 1000 T tTout TTout 900 •4% rms p’/P Cout at turbine inlet 800 700 600 500 0 10 20 30 40 50 5 Time, sec. Pamb P amb 45 4 PCout P Cout 40 Pressure, psia PCCin P CCin 3  P/P cout , % 35 P pj Ppj 30  P/P Cout 2 dP/Pcout 25 1 20 0 15 10 -1 Without Qualification…It Works! 0 10 20 30 40 Time, sec. www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Numerical Investigations What Happens to RPC at Representative P 3 , T 3 ? Approach: •Use in-house 2D axisymmetric CFD code Valve fully closed • Turbulent • Contains detailed chemical kinetics • Adiabatic • Gaseous Jet-A fueled • Successfully applied to PDE, RDE, and SCRAM combustion • Pressure actuated, prescribed motion slide valve simulates reed valve •Validate on atmospheric tests of experimental RPC injector • Compare thrust, mass flow rate, pressure traces, frequency •Run at 10 Atm., 990 R inlet conditions •Optimize for maximum pressure gain at T t4 /T t3 ≈ 2.0 • Fuel injector location • Inlet geometry Valve fully open • Combustion chamber size • Combustor length • Ejector/mixer parameters (length, position, diameter) •Monitor emissions • Seek lowest index with largest pressure gain •Seek minimum size CFD as Predictive Design Tool www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Numerical Investigations Results To Date Inflow Vortex Motion is Key Combustion Chamber: Temperature contours (top half) and fuel Ejector: Length mass fraction contours (bottom half) at various Length Throat simulates NGV b.c. times during one cycle ( � = 0.72). Diameter Throat Diameter Contour Contour Self-ignition via residual hot gas Fuel injection: Rapid confined combustion Placement Timing • Emission Index < 10 g NOX /kg fuel • Lower pressure gain configurations showed Expansion/acceleration values below 1.0! • (P t4 /P t3 - 1)=3.3% @ T t4 /T t3 =2.4 • A large improvement considering T t3 =990 R • Relatively benign station 4 conditions • 7% rms p’/P t4 refill • 23% rms u’/u 4 • 1.7% rms T’/T t4 www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Ongoing and Future Directions Life Extending Techniques Alternative Valve Concepts for Existing Reed Valves Temperature Ejector: Length Throat Diameter Contour Fuel Mass Fraction • Minimum length and diameter configuration • Computational Active Fuel Modulation • Turbine interaction studies • Computational • Active air valves • Still in planning stages • High P 3 , T 3 testing facilities • Still in planning stages AFRL/NASA - 2009 www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration Concluding Remarks Resonant Pulse Combustion (RPC): •Represents a promising approach for achieving practical Pressure Gain Combustion (PGC) •Has features which are well suited for gas turbine applications • Relatively low unsteadiness • Demonstrated approaches to achieving requisite overall lean operation • Few moving parts • Relatively low thermal and mechanical stresses • Self-sustaining • Low emissions potential •Is a remarkably well developed concept •Liquid fueled operation •Demonstrated pressure gain •Demonstrated benefit to gas turbines •Has potential for high P 3 , T 3 operation •Presents multiple opportunities for improvement and optimization that are achievable with current technology RPC Could Be the Gateway to Making PGC Mainstream www.nasa.gov ICVDCW 2017

National Aeronautics and Space Administration END www.nasa.gov ICVDCW 2017