Cavity Flame Holding for High Speed Reacting Flows

2010 ◽  
Author(s):  
Onur Tuncer
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Fuel Injection ◽  
Transport Model ◽  
Leading Edge ◽  
Reacting Flows ◽  
Stagnation Temperature ◽  
Detailed Chemical Kinetics ◽  
Flame Holder ◽  
Shear Stress Transport Model

Combustion phenomena in a ramjet combustor with cavity flame-holder is studied numerically. Combustor follows a constant area isolator and comprises of hydrogen fuel injected sonically upstream of the cavity. Secondary fuel injection is performed at the cavity backwall. A diverging section follows the cavity to prevent thermal choking. These concepts are also utilized in practice. Calculations were performed for an entrance Mach number of 1.4. Stagnation temperature is 702 K, corresponding to a flight Mach number of 3.3 at an altitude of 12.5 km. Detailed chemical kinetics are taken into account with a reaction mechanism comprising of 9 species and 25 reaction steps. Turbulence is modeled using Menter’s k–ω shear stress transport model, which is appropriate for high speed internal flows. It is observed that flame anchors at the leading edge of the cavity, and the flame is stabilized in the cavity mode rather than the jet-wake mode. Numerical simulation captures all the essential features of the reacting flowfield.

Investigation of a Stiff-Integrator Scheme for High-Speed Reacting Flows

2000 ◽  
Author(s):  
Lance D. Woolley ◽  
Douglas A. Schwer ◽  
Russell L. Daines
Keyword(s):  
Time Scales ◽  
High Speed ◽  
Reaction Rates ◽  
Reacting Flows ◽  
Operator Technique ◽  
Detailed Chemical Kinetics ◽  
Time Step ◽  
Kinetics Models ◽  
Split Operator ◽  
Air Chemistry

Abstract Improvements in the modeling of high-speed reacting propulsion flowfields are sought through the coupling of a stiff integrator to determine chemical reaction rates with a multidimensional CFD code. Detailed chemical kinetics models usually have significantly shorter reaction time scales than the fluid time scales, resulting in stiff governing equations and robustness issues. The present work investigates the application of a stiff ordinary differential equation solver, coupled to a diagonalized alternating-direction implicit scheme to decouple the governing time scales. This coupled ODE-ADI split-operator technique is applied to two high-speed reacting flows using hydrogen/air chemistry. The results from the stiff integrator method are compared to the traditional coupled approach utilizing 8- and 18-step kinetics models. Time-step choice, robustness, and comparison of results between the different solution methods are discussed, along with CPU times.

Effect of Centrifugal Force On Gas Flow in a Supersonic Turbine

2021 ◽  
Author(s):  
Takeshi Kanda ◽  
Akio Nakai ◽  
Tatsuya Inagaki ◽  
Tatsuro Asano ◽  
Yasutaka Ohkuma ◽  
...  
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Force ◽  
High Speed ◽  
Gas Flow ◽  
Rocket Engine ◽  
Leading Edge ◽  
Rotor Blades ◽  
Turning Angle ◽  
Liquid Rocket

Abstract The flow condition between the rotor blades of a liquid rocket engine supersonic turbine was studied experimentally and numerically. The entrance Mach number was 1.94, and the turning angle of the blades was 120°. A shock wave was created at the leading edge of the blade, and the Mach number in the passage between the blades decreased to around unity. A similar deceleration has been reported in several past studies. It was found that centrifugal force created the shock wave at the leading edge, reducing both the Mach number and total pressure. This phenomenon is characteristic of high-speed blades with large turning angles. The Mach number in the passage was restricted when the mass flow rate was specified under the specified passage configuration. A convergent-divergent configuration of the passage between the blades suppresses the performance degradation of supersonic turbines.

Experimental investigation of flow topology over a Lambda wing with varying leading edge shape

2018 ◽  
Vol 233 (10) ◽  
pp. 3833-3845
Author(s):  
Liu Changqing ◽  
Li Zengjun ◽  
Liu Qi
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Vortex Formation ◽  
Wind Tunnel Test ◽  
Leading Edge ◽  
Test Model ◽  
Flow Topology ◽  
High Lift ◽  
Oil Flow ◽  
Drag Ratio

In need of high lift/drag ratio design for high-speed unmanned combat aerial vehicles with the Lambda wing configuration, a wind tunnel test is conducted at Mach number 0.4 to select leading edge shapes in consideration of initial compression effect, which in fact takes place during climbout/descending stages. Flow topology over the lee side of the test model is analyzed at Mach number 0.4. Results show that leading edge shapes with a comparatively high lift/drag ratio have a consistent topological structure. Lift/drag properties at the cruise condition are investigated at Mach number 0.75. Aerodynamic forces are measured by a strain gage balance, and the oil flow technique is adopted to obtain flow patterns across the surface of the model. The visualized flow results are further used to analyze vortex formation and development within focused regions.

Computational Study of Mach Number Effects on Dynamic Stall

2019 ◽  
Author(s):  
Taesoon Kim ◽  
Seungtae Kim ◽  
Jiseop Lim ◽  
Junkyu Kim ◽  
Solkeun Jee
Keyword(s):  
Mach Number ◽  
Pressure Gradient ◽  
Rotor Blade ◽  
High Speed ◽  
Computational Study ◽  
Wind Tunnel Test ◽  
Leading Edge ◽  
Adverse Pressure Gradient ◽  
Dynamic Stall ◽  
Navier Stokes

Abstract A rotorcraft main-rotor blade experiences a broad range of the Mach number in high-speed forward flights or rapid maneuvering conditions. Near the boundary of the flight envelope, the rotor blade often encounters severe dynamic stall which limits the overall aerodynamic performance. Of interest here is effects of the Mach number on the dynamic stall. A rotor airfoil, VR-12, is computationally investigated for both static and dynamic stall conditions with varying the Mach number from 0.2 to 0.4. For the small enough Mach number 0.2, both static and dynamic stall are significantly influenced by flow separation due to adverse pressure gradient. For the highest Mach number 0.4, compressibility near the leading edge is no longer negligible, forming a shock there. The shock-induced separation occurs near the leading edge, which dominates both the static and dynamic stall at the high Mach number. At the intermediate Mach number 0.3, it is observed that both adverse pressure gradient and the shock affect the stall. Current computations are conducted with the unsteady Reynolds-averaged Navier Stokes approach with the Spalart-Allmaras model. Numerical results are compared to relevant wind-tunnel test data.

scholarly journals Plasma-Assisted Control of Supersonic Flow over a Compression Ramp

Aerospace ◽  
2019 ◽  
Vol 6 (3) ◽  
pp. 35 ◽  
Author(s):  
Yasumasa Watanabe ◽  
Alec Houpt ◽  
Sergey Leonov
Keyword(s):  
Supersonic Flow ◽  
High Speed ◽  
Leading Edge ◽  
Pressure Change ◽  
Discharge Voltage ◽  
Stagnation Temperature ◽  
Plasma Power ◽  
Supersonic Wind Tunnel ◽  
Compression Ramp ◽  
Plasma Control

This study considers the effect of an electric discharge on the flow structure near a 19.4° compression ramp in Mach-2 supersonic flow. The experiments were conducted in the supersonic wind tunnel SBR-50 at the University of Notre Dame. The stagnation temperature and pressure were varied in a range of 294–600 K and 1–3 bar, respectively, to attain various Reynolds numbers ranging from 5.3 × 105 to 3.4 × 106 based on the distance between the exit of the Mach-2 nozzle and the leading edge of the ramp. Surface pressure measurements, schlieren visualization, discharge voltage and current measurements, and plasma imaging with a high-speed camera were used to evaluate the plasma control authority on the ramp pressure distribution. The plasma being generated in front of the compression ramp shifted the shock position from the ramp corner to the electrode location, forming a flow separation zone ahead of the ramp. It was found that the pressure on the compression surface reduced almost linearly with the plasma power. The ratio of pressure change to flow stagnation pressure was also an increasing function of the ratio of plasma power to enthalpy flux, indicating that the task-related plasma control effectiveness ranged from 17.5 to 25.

Computational Analysis of Hydrogen-Fueled Scramjet Combustor Using Cavities in Tandem Flame Holder

2015 ◽  
Vol 772 ◽  
pp. 130-135 ◽  
Author(s):  
Sukanta Roga ◽  
Krishna Murari Pandey
Keyword(s):  
Mach Number ◽  
Computational Analysis ◽  
Static Pressure ◽  
Combustion Efficiency ◽  
Recirculation Region ◽  
Stagnation Temperature ◽  
Computational Results ◽  
Species Transport ◽  
Flame Holder ◽  
Scramjet Combustor

This work presents the computational analysis of scramjet combustor using cavities in tandem flame holder by means of 3D. The fuel used by scramjet combustor with cavities in tandem flame holder is hydrogen, the fluid flow and the work is based on the species transport combustion with standard k-ε viscous model. The Mach number at inlet is 2.47 and stagnation temperature and static pressure for vitiated air are 1000K and 100kPa respectively. These computational analysis is mainly aimed to study the flow structure and combustion efficiency. The computational results are compared qualitatively and quantitatively with experimental results and these are agreed as well. Due to the combustion, the recirculation region behind the cavity injector becomes larger as compared to mixing case which acts as a flame holder. From the analysis, the maximum Mach number of 2.33 is observed in the recirculation areas.

Stagnation temperature effect on the supersonic flow around pointed airfoils with application for air

2018 ◽  
Vol 19 (3) ◽  
pp. 312
Author(s):  
Rahima Takhnouni ◽  
Toufik Zebbiche ◽  
Abderrazak Allali
Keyword(s):  
High Temperature ◽  
Mach Number ◽  
Supersonic Flow ◽  
Incidence Angle ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Oblique Shock Wave ◽  
Stagnation Temperature ◽  
Perfect Gas ◽  
Application Limit

The aim of this work is to develop a new numerical calculation program to determine the effect of the stagnation temperature on the calculation of the supersonic flow around a pointed airfoils using the equations for oblique shock wave and the Prandtl Meyer expansion, under the model at high temperature, calorically imperfect and thermally perfect gas, lower than the dissociation threshold of the molecules. The specific heat at constant pressure does not remain constant and varies with the temperature. The new model allows making corrections to the perfect gas model designed for low stagnation temperature, low Mach number, low incidence angle and low airfoil thickness. The stagnation temperature is an important parameter in our model. The airfoil should be pointed at the leading edge to allow an attached shock solution to be seen. The airfoil is discretized into several panels on the extrados and the intrados, placed one adjacent to the other. The distribution of the flow on the panel in question gives a compression or an expansion according to the deviation of the flow with respect to the old adjacent panel. The program determines all the aerodynamic characteristics of the flow and in particular the aerodynamic coefficients. The calculation accuracy depends on the number of panels considered on the airfoil. The application is made for high values of stagnation temperature, Mach number and airfoil thickness. A comparison between our high temperature model and the perfect gas model is presented, in order to determine an application limit of the latter. The application is for air.

Preliminary Simulation of a 3D Turbine Stage with In Situ Combustion

2015 ◽  
Vol 772 ◽  
pp. 103-107 ◽  
Author(s):  
Sterian Danaila ◽  
Dragoș Isvoranu ◽  
Constantin Leventiu
Keyword(s):  
Fuel Injection ◽  
Stokes Equations ◽  
Transport Model ◽  
Salient Feature ◽  
Navier Stokes ◽  
Combustion Model ◽  
Turbine Stage ◽  
Shear Stress Transport Model ◽  
Rotor Stator Interaction

This paper presents the preliminary results of the numerical simulation of flow and combustion in a one stage turbine combustor (turbine stage in situ combustion). The main purpose of the simulation is to assess the stability of the in situ combustion with respect to the unsteadiness induced by the rotor-stator interaction. Apart from previous attempts, the salient feature of this CFD approach is the new fuel injection concept that consisting of a perforated pipe placed at mid-pitch in the stator row passage. The flow and combustion are modelled by the Reynolds-averaged Navier-Stokes equations coupled with the species transport equations. The chemistry model used herein is a two-step, global, finite rate combustion model while the turbulence model is the shear stress transport model. The chemistry turbulence interaction is described in terms of eddy dissipation concept.

Unsteady Numerical Simulation of Shock Systems in Vaneless Counter-Rotating Turbine

2005 ◽  
Author(s):  
Huishe Wang ◽  
Qingjun Zhao ◽  
Xiaolu Zhao ◽  
Jianzhong Xu
Keyword(s):  
Numerical Simulation ◽  
Mach Number ◽  
Leading Edge ◽  
Turbine Rotor ◽  
Nozzle Design ◽  
Suction Surface ◽  
Rotor Wake ◽  
Wake Interaction ◽  
Unsteady Numerical Simulation ◽  
Almost All

A detailed unsteady numerical simulation has been carried out to investigate the shock systems in the high pressure (HP) turbine rotor and unsteady shock-wake interaction between coupled blade rows in a 1+1/2 counter-rotating turbine (VCRT). For the VCRT HP rotor, due to the convergent-divergent nozzle design, along almost all the span, fishtail shock systems appear after the trailing edge, where the pitch averaged relative Mach number is exceeding the value of 1.4 and up to 1.5 approximately (except the both endwalls). A group of pressure waves create from the suction surface after about 60% axial chord in the VCRT HP rotor, and those waves interact with the inner-extending shock (IES). IES first impinges on the next HP rotor suction surface and its echo wave is strong enough and cannot be neglected, then the echo wave interacts with the HP rotor wake. Strongly influenced by the HP rotor wake and LP rotor, the HP rotor outer-extending shock (OES) varies periodically when moving from one LP rotor leading edge to the next. In VCRT, the relative Mach numbers in front of IES and OES are not equal, and in front of IES, the maximum relative Mach number is more than 2.0, but in front of OES, the maximum relative Mach number is less than 1.9. Moreover, behind IES and OES, the flow is supersonic. Though the shocks are intensified in VCRT, the loss resulted in by the shocks is acceptable, and the HP rotor using convergent-divergent nozzle design can obtain major benefits.

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