CFD Predicted Discharge Coefficients of a Single Cylindrical Hole With Compressible External Cross-Flow

2009 ◽  
Author(s):  
L. Guo ◽  
Y. Y. Yan ◽  
J. D. Maltson
Keyword(s):  
Mach Number ◽  
Pressure Distribution ◽  
Inclination Angle ◽  
Film Cooling ◽  
Discharge Coefficient ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Cylindrical Hole ◽  
Static Pressure Distribution ◽  
Mainstream Flow

A computational investigation on discharge coefficient (Cd) of a single cylindrical hole is presented in this paper. The numerical calculations are carried out on a 3-D compressible model. The Shear-Stress Transport (SST) k–ω model is used to simulate the turbulence in the flow. The inclination angle (α) of the film cooling hole varies from 20° to 30°, 45° and 90°, respectively. The diameter of the hole is fixed at 10mm, but different coolant to mainstream pressure ratios (ptc/pm) are examined. The coolant Mach number (Mac) is set at a constant value of 0.3 and the mainstream flow Mach number (Mam) varies from 0.3 to 1.4. The effects of Mam and α on the Cd value as well as the static pressure distribution at the jet exit are investigated. The numerical results show an acceptable agreement in the trend of the Cd variation compare with the available experimental data. It has been predicted that the static pressure distribution in the vicinity of the jet exit is influenced by a number of factors including the mainstream flow Mach number, shock wave, jet inclination angle and the pressure ratio of the coolant to the mainstream flow. And then the static pressure field near the hole can further give strong influence on the discharge coefficient.

Discharge Coefficient Measurements of Film-Cooling Holes With Expanded Exits

1998 ◽  
Vol 120 (3) ◽  
pp. 557-563 ◽  
Author(s):  
M. Gritsch ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Discharge Coefficient ◽  
Pressure Ratio ◽  
Cylindrical Hole ◽  
Flow Conditions ◽  
Wide Range ◽  
Cooling Hole ◽  
Shaped Hole ◽  
Film Cooling Holes

This paper presents the discharge coefficients of three film-cooling hole geometries tested over a wide range of flow conditions. The hole geometries include a cylindrical hole and two holes with a diffuser-shaped exit portion (i.e., a fan-shaped and a laidback fan-shaped hole). The flow conditions considered were the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the pressure ratio across the hole (up to 2). The results show that the discharge coefficient for all geometries tested strongly depends on the flow conditions (crossflows at hole inlet and exit, and pressure ratio). The discharge coefficient of both expanded holes was found to be higher than of the cylindrical hole, particularly at low pressure ratios and with a hole entrance side crossflow applied. The effect of the additional layback on the discharge coefficient is negligible.

scholarly journals Discharge Coefficient Measurements of Film-Cooling Holes With Expanded Exits

1997 ◽  
Author(s):  
M. Gritsch ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Discharge Coefficient ◽  
Pressure Ratio ◽  
Cylindrical Hole ◽  
Flow Conditions ◽  
Discharge Coefficients ◽  
Wide Range ◽  
Cooling Hole ◽  
Film Cooling Holes

This paper presents the discharge coefficients of three film-cooling hole geometries tested over a wide range of flow conditions. The hole geometries include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fanshaped and a laidback fanshaped hole). The flow conditions considered were the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the pressure ratio across the hole (up to 2). The results show that the discharge coefficient for all geometries tested strongly depends on the flow conditions (crossflows at hole inlet and exit, and pressure ratio). The discharge coefficient of both expanded holes was found to be higher than of the cylindrical hole, particularly at low pressure ratios and with a hole entrance side crossflow applied. The effect of the additional layback on the discharge coefficient is negligible.

Adjoint Optimization of Multistage Axial Compressor Blades with Static Pressure Constraint at Blade Row Interface

2016 ◽  
Vol 33 (2) ◽  
Author(s):  
Jia Yu ◽  
Lucheng Ji ◽  
Weiwei Li ◽  
Weilin Yi
Keyword(s):  
Pressure Distribution ◽  
Adjoint Method ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Optimal Solution ◽  
Optimization Procedure ◽  
Compressor Blades ◽  
Static Pressure Distribution ◽  
Practical Engineering

AbstractAdjoint method is an important tool for design refinement of multistage compressors. However, the radial static pressure distribution deviates during the optimization procedure and deteriorates the overall performance, producing final designs that are not well suited for realistic engineering applications. In previous development work on multistage turbomachinery blade optimization using adjoint method and thin shear-layer N-S equations, the entropy production is selected as the objective function with given mass flow rate and total pressure ratio as imposed constraints. The radial static pressure distribution at the interfaces between rows is introduced as a new constraint in the present paper. The approach is applied to the redesign of a five-stage axial compressor, and the results obtained with and without the constraint on the radial static pressure distribution at the interfaces between rows are discussed in detail. The results show that the redesign without the radial static pressure distribution constraint (RSPDC) gives an optimal solution that shows deviations on radial static pressure distribution, especially at rotor exit tip region. On the other hand, the redesign with the RSPDC successfully keeps the radial static pressure distribution at the interfaces between rows and make sure that the optimization results are applicable in a practical engineering design.

Shape Optimization of Multi-Stage Axial Compressor Blades Using Adjoint Method With Static Pressure Constraint at Blade Row Interface

2015 ◽  
Author(s):  
Jia Yu ◽  
Lucheng Ji ◽  
Weiwei Li ◽  
Weilin Yi
Keyword(s):  
Pressure Distribution ◽  
Adjoint Method ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Optimal Solution ◽  
Optimization Procedure ◽  
Compressor Blades ◽  
Static Pressure Distribution ◽  
Multi Stage

Adjoint method is an important tool for design refinement of multistage compressors. However, the radial static pressure distribution deviates during the optimization procedure and deteriorates the overall performance, producing final designs that are not well suited for realistic engineering applications. In previous development work on multistage turbomachinery blade optimization using adjoint method and thin shear-layer N-S equations, the entropy production is selected as the objective function with given mass flow rate and total pressure ratio as imposed constraints. The radial static pressure distribution at the interfaces between rows is introduced as a new constraint in the present paper. The approach is applied to the redesign of a five-stage axial compressor, and the results obtained with and without the constraint on the radial static pressure distribution at the interfaces between rows are discussed in detail. The results show that the redesign without radial static pressure distribution constraint (RSPDC) gives an optimal solution that shows deviations on radial static pressure distribution, especially at rotor exit tip region. On the other hand, the redesign with the RSPDC successfully keeps the radial static pressure distribution at the interfaces between rows and make sure that the optimization results are applicable in a practical engineering design.

scholarly journals THE EFFECT OF CONVERGENT-DIVERGENT NOZZLE PROFILE ON ITS PERFORMANCE

2019 ◽  
Vol 19 (1) ◽  
pp. 14-43
Author(s):  
Arkan Al-Taie ◽  
Hussien W Mashi ◽  
Ali M Hadi
Keyword(s):  
Mach Number ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Sharp Decrease ◽  
Inlet Pressure ◽  
Ansys Fluent ◽  
Static Pressure Distribution ◽  
Flow Parameters ◽  
Sudden Rise ◽  
Flow Through

The paper presents the effect of convergent-divergent nozzles profile across specified inlet pressures values from (1.5 bar-4 bar), with constant back pressure of (1 bar). The flow of air through three convergent-divergent nozzles was studied theoretically. The flow was assumed to be one-dimensional, adiabatic and reversible (isentropic). The flow parameters like static pressure ratio and Mach number were analyzed. The flow parameters were obtained in term of area ratio along the nozzle. MATLAB code was built in order to find the Mach number along the nozzles, by using Newton-Raphson method. The shockwave position inside the nozzles was determined, using "analytic method". ANSYS fluent 18 was used to simulate the flow through the three nozzles. Two- dimensional, turbulent and viscous models were utilized to solve the governing equations. K-? model was used to model the turbulent effect. The results concluded that, reduction in inlet pressure can not affect the flow upstream the throat. Also the shockwave appearance can be noticed by a sudden rise in static pressure associated with a sharp decrease in Mach number. Shockwave moves toward the throat by reduction the inlet total pressure .By comparison the static pressure distribution along the three nozzles where can be deduced that the profile has an effect on the flow character i.e. (static pressure Mach no).The best performance among the nozzles is the performance of nozzle (N1), which (75%) of its length work as nozzle at the lowest inlet pressure of (1.5bar) while (44% and 60%) of the nozzles length for (N2 and N3) respectively work as the nozzle.

Ingestion Into the Upstream Wheelspace of an Axial Turbine Stage

1994 ◽  
Vol 116 (2) ◽  
pp. 327-332 ◽  
Author(s):  
T. Green ◽  
A. B. Turner
Keyword(s):  
Pressure Distribution ◽  
Static Pressure ◽  
Three Dimensional ◽  
Computer Code ◽  
Rotor Blades ◽  
Rotor Speed ◽  
Turbine Stage ◽  
Flow Rates ◽  
Static Pressure Distribution ◽  
Axial Clearance

The upstream wheelspace of an axial air turbine stage complete with nozzle guide vanes (NGVs) and rotor blades (430 mm mean diameter) has been tested with the objective of examining the combined effect of NGVs and rotor blades on the level of mainstream ingestion for different seal flow rates. A simple axial clearance seal was used with the rotor spun up to 6650 rpm by drawing air through it from atmospheric pressure with a large centrifugal compressor. The effect of rotational speed was examined for several constant mainstream flow rates by controlling the rotor speed with an air brake. The circumferential variation in hub static pressure was measured at the trailing edge of the NGVs upstream of the seal gap and was found to affect ingestion significantly. The hub static pressure distribution on the rotor blade leading edges was rotor speed dependent and could not be measured in the experiments. The Denton three-dimensional C.F.D. computer code was used to predict the smoothed time-dependent pressure field for the rotor together with the pressure distribution downstream of the NGVs. The level and distribution of mainstream ingestion, and thus the seal effectiveness, was determined from nitrous oxide gas concentration measurements and related to static pressure measurements made throughout the wheelspace. With the axial clearance rim seal close to the rotor the presence of the blades had a complex effect. Rotor blades in connection with NGVs were found to reduce mainstream ingestion seal flow rates significantly, but a small level of ingestion existed even for very high levels of seal flow rate.

An Investigation of Flow Characteristics and Parameter Effects for a New Concept of Hybrid SVC Nozzle

2018 ◽  
Author(s):  
F. Song ◽  
J. W. Shi ◽  
L. Zhou ◽  
Z. X. Wang ◽  
X. B. Zhang
Keyword(s):  
Static Pressure ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Exhaust System ◽  
Aircraft Engine ◽  
Thrust Coefficient ◽  
Thrust Vectoring ◽  
Static Pressure Distribution ◽  
Thrust Vector ◽  
Vector Angle

Lighter weight, simpler structure, higher vectoring efficiency and faster vector response are recent trends in development of aircraft engine exhaust system. To meet these new challenges, a concept of hybrid SVC nozzle was proposed in this work to achieve thrust vectoring by adopting a rotatable valve and by introducing a secondary flow injection. In this paper, we numerically investigated the flow mechanism of the hybrid SVC nozzle. Nozzle performance (e.g. the thrust vector angle and the thrust coefficient) was studied with consideration of the influence of aerodynamic and geometric parameters, such as the nozzle pressure ratio (NPR), the secondary pressure ratio (SPR) and the deflection angle of the rotatable valve (θ). The numerical results indicate that the introductions of the rotatable valve and the secondary injection induce an asymmetrically distributed static pressure to nozzle internal walls. Such static pressure distribution generates a side force on the primary flow, thereby achieving thrust vectoring. Both the thrust vector angle and vectoring efficiency can be enhanced by reducing NPR or by increasing θ. A maximum vector angle of 16.7 ° is attained while NPR is 3 and the corresponding vectoring efficiency is 6.33 °/%. The vector angle first increases and then decreases along with the elevation of SPR, and there exists an optimum value of SPR for maximum thrust vector angle. The effects of θ and SPR on the thrust coefficient were found to be insignificant. The rotatable valve can be utilized to improve vectoring efficiency and to control the vector angle as expected.

scholarly journals Adiabatic Wall Effectiveness Measurements of Film-Cooling Holes With Expanded Exits

1997 ◽  
Author(s):  
M. Gritsch ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Thermal Protection ◽  
Lateral Spreading ◽  
Cylindrical Hole ◽  
Adiabatic Wall ◽  
Camera System ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole

This paper presents detailed measurements of the film-cooling effectiveness for three single, scaled-up film-cooling hole geometries. The hole geometries investigated include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fanshaped and a laidback fanshaped hole). The flow conditions considered are the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the blowing ratio (up to 2). The coolant-to-mainflow temperature ratio is kept constant at 0.54. The measurements are performed by means of an infrared camera system which provides a two-dimensional distribution of the film-cooling effectiveness in the nearfield of the cooling hole down to x/D = 10. As compared to the cylindrical hole, both expanded holes show significantly improved thermal protection of the surface downstream of the ejection location, particularly at high blowing ratios. The laidback fanshaped hole provides a better lateral spreading of the ejected coolant than the fanshaped hole which leads to higher laterally averaged film-cooling effectiveness. Coolant passage crossflow Mach number and orientation strongly affect the flowfield of the jet being ejected from the hole and, therefore, have an important impact on film-cooling performance.

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