High Fidelity Modeling and Simulation of Plasma Assisted Combustion in Complex Flow Environments Vigor Yang Daniel Guggenheim School of Aerospace Engineering Georgia Institute of Technology Atlanta, Georgia 30332 MURI Kick-Off Meeting Ohio State University, Columbus, Ohio November 4, 2009
Work Scope School of Aerospace Engineering Establish an integrated theoretical and numerical framework to substantially improve understanding of plasma-assisted ignition and combustion in complex flow environments typical of flight engines. -- steady and transient operating conditions -- entire flowpath -- plasma-activated flow control Focus on -- laminar flow experiments (Task 1) -- ignition and flameholding experiments for nonequilibrium plasma flows (Tasks 12 and 13) 2
AFRL scramjet combustion test facility (test cell 18) nozzle isolator combustor I-1 I-2 throttle 1.23 cm I-3 I-4 aerothrottle 1.27 cm 2.54 cm 2.54 cm flow spark slug pilot fuel z 2.54 cm 2.54 cm 5.08 cm I-1 I-2 1.27 cm Body side Cowl side Research Objectives to establish a unified 3-D theoretical/numerical framework for treating unsteady flow and flame dynamics in scramjet propulsion p systems to investigate the effects of air throttling on ignition enhancement and flameholding mechanisms in a scramjet combustor 3
Computational Domain School of Aerospace Engineering Flight conditions: flight Mach number: 3.5 5-5 5 flight altitude: 15-28 km dynamic pressure: 24-96 kpa T 0 = 1106.7 K p 0 = 3.51 atm Facility nozzle: Mach 2.2 Cavity configuration: length-to-depth ratio: L/D = 6.1 Computational domain: (1654x95+361x85)x173 = 32.5 million grids 361x85x173 = 5.3 million grids in cavity 164 numerical blocks 4
AFRL Scramjet Test Cell 18 Facility operation conditions code validation M q(kpa) T 0 (k) p 0 (atm) m air (kg/s) 5.0 96 1062.8 14.2 3.117 5.0 48 1083.3 7.0 1.524 5.0 24 1106.7 3.5 0.757 4.5 96 981.1 14.4 3.288 4.5 48 995.5 7.0 1.601 4.5 24 1015.5 3.5 0.789 5
Ignition Operation Sequence School of Aerospace Engineering ignition on ignition off air throttli ttling off Ignition operation sequence cold flow starts fuel inject ection on air throttl ttling on with throttling t 0 =0 t 1 t 2 t 3 t 4 without throttling flame stabilized t 5 flame stabilized or blowout fuel inje njection on cold flow starts ignition n off ignition on Engine operation conditions Case M in U in p in T in φ m air m throttle 1 2.2 1045 0328 560 0.6 0.757 0.0 2 2.2 1045 0.328 560 0.6 0.757 0.151 6
Shadowgraph of Cold Flowfield Baseline Study shadowgraph of cold flowfield (z/w = 1/2) combustor section expansion wave emanating from the front edge of the cavity free shear layer attached at the tail of cavity strong reattachment shock originating from the rear part of the cavity 7
Shadowgraph of Cold Flowfield Baseline Study shadowgraph of cold flowfield (z/w = 1/2) isolator section 8
Steady State of Cold Flow with Air Throttling 3-D perspective shadowgraph view shear layer impingement on the rear edge of the cavity eventually lifted into the downstream separated boundary layer a series of oblique shock waves reflection by means of the shock/shear layer interactions bow-shaped shock structure at orifices 9
Effect of Air Throttling on Cold Flow Development shadowgraph view of cold flow evolution wall pressure evolution throttle injector: 3-section slit sonic airflow normal to the main air flow operation conditions: throttle turned on at t = 0.0 ms and x = 1.36 m throttling mass flow rate: 0.15 kg/s, 20 % m air 10
Effect of Air Throttling on Cold Flow Development shadowgraph view of cold flow evolution wall pressure evolution throttle injector: 3-section slit sonic airflow normal to the main air flow operation conditions: throttle turned on at t = 0.0 ms and x = 1.36 m throttling mass flow rate: 0.151 kg/s, 20 % m air 11
Steady State of Cold Flow with Air Throttling School of Aerospace Engineering vertical shadowgraph view horizontal shadowgraph view shear layer impingement on the rear edge of the cavity eventually lifted reflection of a series of oblique shock inside combustor induces shock/boundary layer interaction bow-shaped shockstructure shocks at orifices 12
Mass Fraction Contours of Ethylene (%) no air throttling air throttling 13
Steady State of Cold Flow with Air Throttling perspective vorticity magnitude and streamlines shock/boundary-layer interaction enhances flow distortion and turbulent diffusion in the combustor, accompanied with significant vorticity magnitude intense large recirculation zones were generated inside id cavity upstream of air throttling 14
Ignition Transient & Flame Development 15
Ignition Transient and Flame Evolution (no air throttling) School of Aerospace Engineering Corresponding flight conditions: T 0 = 1106.7 K, p 0 = 3.5 atm q = 24 kpa flight Mach number: 5 flight altitude: 30 km Ignition Transient without Air Throttling t = 0.0 ms, igniter turned on t = 1.0 ms, fuel/air mixture ignited in the cavity, and igniter turned off t ~ 2.0 ms, flame spreads into the chamber, and blown off by the main stream t = 5.0 ms, flame disappears in cavity, no combustion observed at cowl-side walls and chamber 16
Ignition Transient and Flow Structure at t = 3.563 ms (no air throttling) School of Aerospace Engineering shadowgraph image flame wall pressure and Mach number reacting flow anchored in the cavity region weak combustion in cavity weak sidewall and corner effects 17
flame evolution Ignition Transient and Flame Evolution (air throttling) operation conditions: igniter located at bottom of cavity igniter power < 100 w fuel mass flow rate: m fuel = 0.048 kg/s fuel/air total equivalence ratio: = 0.6 throttling mass flow rate: 0.15 kg/s, 20 % m air Operation sequence: t = 0.0 ms, igniter turned on t = 0.7 ms, flame established in cavity, and igniter turned off t = 4.8 ms, flame established on both side-walls in combustor, and throttle turned off t = 9.6 ms, flame sustained and stabilized in combustor 18
Effect of Air Throttling on Reacting Flow Development 19
Formation of Pre-combustion Shock Train (t = 4.811 ms) near the vertical sidewall on the centre plane wall pressure and Mach number boundary layer separation at corners due to adverse pressure gradient primary formation of pre-combustion shock train in isolator at t = 4.811 ms main flow speed reduced increasing of combustor pressure by heat release 20
Pre-combustion Shock Train in Stable Reacting Flow shadowgraph image on the centre plane oblique shock train length x Waltrup and Billig (1973) developed d a correlation for the length of the shock train using axisymmetric ducts with downstream flow throttling to generate shock trains 2 x M 2 1/5 1 1Re 50 p f f 1 170 p predicted length of shock train: 27.5 cm 1 1/2 1/2 H pa pa measured dlength of shock ktrain: 32 cm Model input conditions * M 1 H(m) θ(mm) R eθ p f (atm) p a (atm ) 1.9 0.038 9.0 8750 1.4 0.4 * x: the length of shock train M 1 : Mach number of approach flow H: duct height θ: boundary layer momentum thickness for undisturbed flow R eθ : Reynolds number based on boundary layer momentum thickness p f /p a : ratio of the local wall pressure to the static pressure at the beginning of the pressure rise 21
Ignition Transient on Cowl-side Wall (air throttling) Corresponding flight conditions: T 0 = 1106.7 K, p 0 = 3.5 atm q = 24 kpa flight Mach number: 5 flight altitude: 30 km flame establishes on cowl-side wall, and grows up high temperature region forms center ethylene jet ignited by flame 22
Flame Evolution in Combustor (air throttling) School of Aerospace Engineering Corresponding flight conditions: T 0 = 1106.7 K, p 0 = 3.5 atm q = 24 kpa flight Mach number: 5 flight altitude: 30 km Operation sequence: t = 0.0 ms, igniter turned on t = 0.7 ms, flame established in cavity, and igniter turned off t = 4.8 ms, flame established on both side-walls in combustor, and throttle turned off t = 9.6 ms, flame sustained and stabilized in combustor 23
Ignition Transient on Cowl-side Wall (air throttling) 24
Ignition Transient on Cowl-side Wall (air throttling) Corresponding flight conditions: T 0 = 1106.7 K, p 0 = 3.5 atm q = 24 kpa flight Mach number: 5 flight altitude: 30 km flame established in cavity, and spreads downstream into chamber ignition of ethylene jet on cowl wall supersonic flow core region hot reaction product spreads down along side walls 25
Ignition Transient on Cowl-side Wall (air throttling) cross-section temperature distribution cross-section axial velocity distribution cross-section velocity vector combustion established on the body-side wall low-momentum flow recirculation at the corner regions expansion and sustaining of the flame at the corner regions boundary layer momentum significantly reduced near the sidewalls residence time increased auto-ignition of ethylene on cowl surface 26
cross-section temperature distribution Ignition Transient on Cowl-side Wall (air throttling) cross-section axial velocity distribution two combustion zones in cavity and on cowl-side wall ignited reacting flow propagate downstream, and spread as the second ignition of the fuel jet flow in the center region low-momentum flow recirculation formed inside cavity boundary layer momentum significantly reduced near the cowl-side wall residence time increased flame survived from main stream 27
Stable Flame Structure in Combustor School of Aerospace Engineering Corresponding flight conditions: T 0 = 1106.7 K, p 0 = 3.5 atm q = 24 kpa flight Mach number: 5 flight altitude: 30 km flame established on both walls, and sustained in combustor recirculation zone generated and enhanced by pre-combustion shock train upstream of combustor second flame-holding region forms at corner between side and bottom wall 28
Stable State Flow Structure in Combustor School of Aerospace Engineering shadowgraph in cavity region vorticity magnitude distribution shock wave moves and interacts with boundary layer new vertical structures causes large mount of fluctuation of axial velocity into combustor add significant ifi vorticity it in shear layer, roll up into large-scale l vortices across the cavity mixing enhanced due to increasing area contact between low- and high-momentum flow pre-combustion shock train upstream lead to significant influence to stabilization of combustion 29
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