Two and three dimensional parabolized Navier-Stokes code for scramjet combustor, nozzle, and film cooling analysis

1992 ◽  
Author(s):  
HOUSHANG EBRAHIMI ◽  
MIKE GILBERTSON
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Scramjet Combustor

Film Cooling From Novel Sister Shaped Single-Holes

2014 ◽  
Author(s):  
Siavash Khajehhasani ◽  
Bassam Jubran
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Lateral Spreading ◽  
Navier Stokes ◽  
The Novel ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Shaped Hole ◽  
Lift Off

A numerical investigation of the film cooling performance from novel sister shaped single-holes (SSSH) is presented in this paper and the obtained results are compared with a single cylindrical hole, a forward diffused shaped hole, as well as discrete sister holes. Three types of the novel sister shaped single-hole schemes namely downstream, upstream and up/downstream SSSH, are designed based on merging the discrete sister holes to the primary hole in order to reduce the jet lift-off effect and increase the lateral spreading of the coolant on the blade surface as well as a reduction in the amount of coolant in comparison with discrete sister holes. The simulations are performed using three-dimensional Reynolds-Averaged Navier Stokes analysis with the realizable k–ε model combined with the standard wall function. The upstream SSSH demonstrates similar film cooling performance to that of the forward diffused shaped hole for the low blowing ratio of 0.5. While it performs more efficiently at M = 1, where the centerline and laterally averaged effectiveness results improved by 70% and 17%, respectively. On the other hand, the downstream and up/downstream SSSH schemes show a considerable improvement in film cooling performance in terms of obtaining higher film cooling effectiveness and less jet lift-off effect as compared with the single cylindrical and forward diffused shaped holes for both blowing ratios of M = 0.5 and 1. For example, the laterally averaged effectiveness for the downstream SSSH configuration shows an improvement of approximately 57% and 110% on average as compared to the forward diffused shaped hole for blowing ratios of 0.5 and 1, respectively.

A CFD-Based Parametric Study on Modification of Discrete Sister Holes Location for the Film Cooling Flow

2014 ◽  
Author(s):  
Siavash Khajehhasani ◽  
Bassam Jubran
Keyword(s):  
Film Cooling ◽  
Numerical Study ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Base Case ◽  
Spanwise Direction ◽  
Cooling Performance ◽  
Cooling Flow ◽  
Lift Off ◽  
The Individual

A numerical study on the effects of sister holes locations on film cooling performance is presented. This includes the change of the location of the individual discrete sister holes in the streamwise and spanwise directions, where each one of these directions includes 9 different locations, The simulations are performed using three-dimensional Reynolds-Averaged Navier Stokes analysis with the realizable k–ε model combined with the standard wall function. The variation of the sister holes in the streamwise direction provides similar film cooling performance as the base case for both blowing ratios of 0.5 and 1. On the other hand, the spanwise variation of the sister holes’ location has a more prominent effect on the effectiveness. In some cases, as a result of the anti-vortices generated from the sister holes and the repositioning of the sister holes in the spanwise direction, the jet lift-off effect notably decreases and more volume of coolant is distributed in the spanwise direction.

scholarly journals Heat Transfer on a Film-Cooled Rotating Blade

1999 ◽  
Author(s):  
Vijay K. Garg
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Three Dimensional ◽  
Leading Edge ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

A multi-block, three-dimensional Navier-Stokes code has been used to compute heat transfer coefficient on the blade, hub and shroud for a rotating high-pressure turbine blade with 172 film-cooling holes in eight rows. Film cooling effectiveness is also computed on the adiabatic blade. Wilcox’s k-ω model is used for modeling the turbulence. Of the eight rows of holes, three are staggered on the shower-head with compound-angled holes. With so many holes on the blade it was somewhat of a challenge to get a good quality grid on and around the blade and in the tip clearance region. The final multi-block grid consists of 4784 elementary blocks which were merged into 276 super blocks. The viscous grid has over 2.2 million cells. Each hole exit, in its true oval shape, has 80 cells within it so that coolant velocity, temperature, k and ω distributions can be specified at these hole exits. It is found that for the given parameters, heat transfer coefficient on the cooled, isothermal blade is highest in the leading edge region and in the tip region. Also, the effectiveness over the cooled, adiabatic blade is the lowest in these regions. Results for an uncooled blade are also shown, providing a direct comparison with those for the cooled blade. Also, the heat transfer coefficient is much higher on the shroud as compared to that on the hub for both the cooled and the uncooled cases.

scholarly journals Numerical Study on the Solid Fuel Rocket Scramjet Combustor with Cavity

Energies ◽  
2019 ◽  
Vol 12 (7) ◽  
pp. 1235 ◽  
Author(s):  
Chaolong Li ◽  
Zhixun Xia ◽  
Likun Ma ◽  
Xiang Zhao ◽  
Binbin Chen
Keyword(s):  
Solid Fuel ◽  
Stokes Equations ◽  
Numerical Study ◽  
Three Dimensional ◽  
Combustion Efficiency ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Particle Combustion ◽  
Scramjet Combustor ◽  
Cavity Configuration

Scramjet based on solid propellant is a good supplement for the power device of future hypersonic vehicles. A new scramjet combustor configuration using solid fuel, namely, the solid fuel rocket scramjet (SFRSCRJ) combustor is proposed. The numerical study was conducted to simulate a flight environment of Mach 6 at a 25 km altitude. Three-dimensional Reynolds-averaged Navier–Stokes equations coupled with shear stress transport (SST) k − ω turbulence model are used to analyze the effects of the cavity and its position on the combustor. The feasibility of the SFRSCRJ combustor with cavity is demonstrated based on the validation of the numerical method. Results show that the scramjet combustor configuration with a backward-facing step can resist high pressure generated by the combustion in the supersonic combustor. The total combustion efficiency of the SFRSCRJ combustor mainly depends on the combustion of particles in the fuel-rich gas. A proper combustion organization can promote particle combustion and improve the total combustion efficiency. Among the four configurations considered, the combustion efficiency of the mid-cavity configuration is the highest, up to about 70%. Therefore, the cavity can effectively increase the combustion efficiency of the SFRSCRJ combustor.

scholarly journals Computations of Film Cooling From Holes With Struts

1999 ◽  
Author(s):  
T. I-P. Shih ◽  
Y.-L. Lin ◽  
M. K. Chyu ◽  
S. Gogineni
Keyword(s):  
Film Cooling ◽  
Computational Study ◽  
Three Dimensional ◽  
Protective Layer ◽  
Navier Stokes ◽  
Time Stepping ◽  
Cooling Hole ◽  
A Cell ◽  
Volume Method ◽  
Film Cooling Holes

Computations were performed to study the three-dimensional flow and heat transfer on a flat plate cooled by jets, injected from a plenum through one row of film-cooling holes in which each hole is fitted with a strut in the form of a circular cylinder. Three different configurations of the film-cooling hole were investigated: without strut, with streamwise strut, and with spanwise strut. For all configurations, the film-cooling holes are inclined at 35°, and the coolant-to-mainflow density and mass-flux ratios are 1.6 and 0.5, respectively. The focus of this study is to understand how struts in holes affect film cooling jets and their interactions with the mainflow in forming a protective layer of cooler fluid over the plate. This computational study is based on the ensemble-averaged conservation equations of mass, momentum (compressible Navier-Stokes), and energy. Effects of turbulence was modeled by a low Reynolds number k-ω closure known as the shear-stress-transport (SST) model. Solutions were generated by a cell-centered finite-volume method that uses third-order accurate flux-difference splitting of Roe with limiters, multigrid acceleration of a diagonalized ADI scheme with local time stepping, and patched/overlapped structured grids. In the computations, the flow is resolved not just in the cooling-jet/mainflow interaction region, but also inside the film-cooling holes and in the plenum. Computed results for adiabatic effectiveness for the case without struts were compared with experimental data, and reasonably good agreements were obtained.

Investigation of Discrete-Hole Film Cooling Parameters Using Curved-Plate Models

1999 ◽  
Vol 121 (4) ◽  
pp. 792-803 ◽  
Author(s):  
M. K. Berthe ◽  
S. V. Patankar
Keyword(s):  
Film Cooling ◽  
Stokes Equations ◽  
Turbulent Viscosity ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Grid Turbulence ◽  
Injection Angle ◽  
Flat Surfaces ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

Computations have been conducted on curved, three-dimensional discrete-hole film cooling geometries that included the mainflow, injection hole, and supply plenum regions. Both convex and concave film cooling geometries were studied. The effects of several film cooling parameters have been investigated, including the effects of blowing ratio, injection angle, hole length, hole spacing, and hole staggering. The blowing ratio was varied from 0.5 to 1.5, the injection angle from 35 to 65 deg, the hole length from 1.75D to 6.0D, and the hole spacing from 2D to 3D. The staggered-hole arrangement considered included two rows. The computations were performed by solving the fully elliptic, three-dimensional Navier–Stokes equations over a body-fitted grid. Turbulence closure was achieved using a modified k–ε model in which algebraic relations were used for the turbulent viscosity and the turbulent Prandtl number. The results presented and discussed include plots of adiabatic effectiveness as well as plots of velocity contours and velocity vectors in cross-stream planes. The present study reveals that the blowing ratio, hole spacing, and hole staggering are among the most significant film cooling parameters. Furthermore: (1) The optimum blowing ratios for curved surfaces are higher than those for flat surfaces, (2) a reduction of hole spacing from 3D to 2D resulted in a very significant increase in adiabatic effectiveness, especially on the concave surface, (3) the increase in cooling effectiveness with decreasing hole spacing was found to be due to not only the increased coolant mass per unit area, but also the smaller jet penetration and the weaker counterrotating vortices, (4) for all practical purposes, the hole length was found to be a much less significant film cooling parameter.

Analyzes of Film Cooling Effectiveness from Cylindrical and Staggered Compound Cooling Holes with Alignment Angle of 30 Degree at the End of Combustor

2014 ◽  
Vol 695 ◽  
pp. 389-392
Author(s):  
Shahin Salimi ◽  
Nor Azwadi Che Sidik ◽  
Leila Jahanshaloo ◽  
Kianpour Ehsan
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Internal Cooling ◽  
Ansys Fluent ◽  
Cooling Effectiveness ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Alignment Angle ◽  
Compound Angle

A numerical simulation has been performed for the investigation of flow and heat transfer characteristics of a film cooling injected through a hole with cylindrical and compound angle orientation. This paper presents the effects of coolant injector configuration of cylindrical and compound cooling holes with alignment angle of 30 degree at blowing ratio, BR = 3.18 on the film cooling effectiveness near the end wall surface of a combustor simulator. In the current research a three dimensional representation of Pratt and Whitney gas turbine engine was simulated and analyzed with a commercial finite volume package ANSYS FLUENT 14.0. This study has been performed with Reynolds-averaged Navier-Stokes turbulence model (RANS) on internal cooling passages The results indicate that using compound angle cooling holes injection, give much better protection than that obtained when simple angle cooling holes were used.

Computational Investigation of Film Cooling from Cylindrical and Row Trenched Cooling Holes near the Combustor End Wall

2014 ◽  
Vol 554 ◽  
pp. 225-229 ◽  
Author(s):  
Nor Azwadi Che Sidik ◽  
Kianpour Ehsan
Keyword(s):  
Film Cooling ◽  
Surface Film ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Internal Cooling ◽  
Cooling Performance ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Three Dimensional Representation

This study was accomplished in order to investigate the effects of cylindrical and row trenched cooling holes with alignment angle of 0 degree and 90 degree at blowing ratio, BR = 3.18 on the film cooling performance adjacent to the endwall surface of a combustor simulator. In this research a three dimensional representation of Pratt and Whitney gas turbine engine was simulated and analyzed with a commercial finite volume package FLUENT 6.2. The current study has been performed with Reynolds-averaged Navier-Stokes turbulence model (RANS) on internal cooling passages. This combustor simulator combined the interaction of two rows of dilution jets, which were staggered in the stream wise direction and aligned in the span wise direction, with that of film cooling along the combustor liner walls. The findings of the study declared that with using the row trenched holes near the endwall surface, film cooling effectiveness is doubled compared to the cooling performance of baseline case.

Effect of Various Injection Hole Shapes on Film Cooling of Turbine Blade

2003 ◽  
Author(s):  
S.-M. Kim ◽  
Youn J. Kim
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Three Dimensional ◽  
Cylindrical Body ◽  
Navier Stokes ◽  
Cylinder Surface ◽  
Injection Hole ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Cooling Flow

Dispersion of coolant jets in a film cooling flow field is the result of a highly complex interaction between the film cooling jets and the mainstream. In order to investigate the effects of injection hole shapes and injection angle on the film cooling of turbine blade, four models having cylindrical and laterally-diffused holes were used. Three-dimensional Navier-Stokes code with k – ε model was used to compute the film cooling coefficient on the turbine blade. A multi-block grid system was generated that was nearly orthogonal to the various surfaces. Mainstream Reynolds number based on the cylinder diameter was 7.1 × 104. The turbulence intensity kept at 5.0% for all inlets. The effect of coolant flow rates was studied for blowing ratios of 0.9, 1.3 and 1.6, respectively. The temperature distribution of the cylindrical body surface is visualized by infrared thermography (IRT) and compared with computational results. Results show that the effects of injection hole shape and injection angle increase as the blowing ratio increases. As lateral injection angle increases, the adiabatic film cooling effectiveness is more broadly distributed and the area protected by coolant increases. The mass flow rate of the coolant through the first-row holes is less than that through the second-row holes due to the pressure distribution around the cylinder surface.

Heat Transfer in Film-Cooled Turbine Blades

1993 ◽  
Author(s):  
Vijay K. Garg ◽  
Raymond E. Gaugler
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Film Cooling ◽  
Stokes Equations ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Explicit Finite Difference ◽  
Control Volumes

In order to study the effect of film cooling on the flow and heat transfer characteristics of actual turbine blades, a three-dimensional Navier-Stokes code has been developed. An existing code (Chima and Yokota, 1990) has been modified for the purpose. The code is an explicit finite difference code with an algebraic turbulence model. The thin-layer Navier-Stokes equations are solved using a general body-fitted coordinate system. The effects of film cooling have been incorporated into the code in the form of appropriate boundary conditions at the hole locations on the blade surface. Each hole exit is represented by several control volumes, thus providing an ability to study the effect of hole shape on the film-cooling characteristics. Comparison with experimental data is fair. Further validation of the code is required, however, and in this respect, there is an urgent need for detailed experimental data on actual turbine blades.

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