Discharge Coefficients of Critical Venturi Nozzles for CO2 and SF6

2000 ◽  
Vol 122 (4) ◽  
pp. 730-734 ◽  
Author(s):  
Shin-ichi Nakao ◽  
Masaki Takamoto
Keyword(s):  
Boundary Layer ◽  
Large Deviation ◽  
Nonequilibrium Flow ◽  
Theoretical Estimation ◽  
Perfect Gas ◽  
Number Range ◽  
Real Gas ◽  
Discharge Coefficients ◽  
Isentropic Flow ◽  
Real Gas Effects

The discharge coefficients of critical Venturi nozzles were measured for CO2 and SF6 on the Reynolds number range from 3×103 to 2×105. The results showed that the measured discharge coefficients for both gases were about 2 percent larger than the theoretical estimation based on the assumption of isentropic flow of a perfect gas and this large deviation could not be reduced even by introducing real gas effects. The experimental results also showed that the large deviation for CO2 could be explained through the assumption of a nonequilibrium flow at the throat. On the other hand, the reason of the deviation observed for SF6 has not been clear yet, but one possible explanation would be the inadequate estimation of the boundary layer at the throat because the theory is based on the laminar boundary layer of a perfect gas. [S0098-2202(00)02004-6]

The interaction of a shock wave with a laminar boundary layer at a compression corner in high-enthalpy flows including real gas effects

1997 ◽  
Vol 342 ◽  
pp. 1-35 ◽  
Author(s):  
S. G. MALLINSON ◽  
S. L. GAI ◽  
N. R. MUDFORD
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Shock Wave ◽  
Gas Flows ◽  
Perfect Gas ◽  
Experimental Conditions ◽  
Real Gas ◽  
High Enthalpy ◽  
Compression Corner ◽  
Real Gas Effects

The high-enthalpy, hypersonic flow over a compression corner has been examined experimentally and theoretically. Surface static pressure and heat transfer distributions, along with some flow visualization data, were obtained in a free-piston shock tunnel operating at enthalpies ranging from 3 MJ kg−1 to 19 MJ kg−1, with the Mach number varying from 7.5 to 9.0 and the Reynolds number based on upstream fetch from 2.7×104 to 2.7×105. The flow was laminar throughout. The experimental data compared well with theories valid for perfect gas flow and with other relevant low-to-moderate enthalpy data, suggesting that for the current experimental conditions, the real gas effects on shock wave/boundary layer interaction are negligible. The flat-plate similarity theory has been extended to include equilibrium real gas effects. While this theory is not applicable to the current experimental conditions, it has been employed here to determine the potential maximum effect of real gas behaviour. For the flat plate, only small differences between perfect gas and equilibrium gas flows are predicted, consistent with experimental observations. For the compression corner, a more rapid rise to the maximum pressure and heat transfer on the ramp face is predicted in the real gas flows, with the pressure lying slightly below, and the heat transfer slightly above, the perfect gas prediction. The increase in peak heat transfer is attributed to the reduction in boundary layer displacement thickness due to real gas effects.

Performance and optimization of gas turbines with real gas effects

2001 ◽  
Vol 215 (4) ◽  
pp. 507-512 ◽  
Author(s):  
A Guha
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Gas Analysis ◽  
Perfect Gas ◽  
Explicit Equation ◽  
Optimum Pressure ◽  
Real Gas ◽  
Turbine Performance ◽  
Real Gas Effects ◽  
Optimum Relation

The influence of various levels of mathematical modelling on gas turbine performance is systematically analysed. It is shown that internal combustion with real gas effects gives rise to an optimum turbine entry temperature which does not arise in a perfect gas analysis and has not been described previously in the literature. At any pressure ratio, the maximum possible efficiency with real gas effects is significantly lower (15-20 per cent) than the maximum possible value predicted by a perfect gas analysis. An explicit equation has been derived for determining the optimum pressure ratio as a function of turbine entry temperature and component efficiencies. It is shown that the optimum design depends very strongly on turbine and compressor efficiencies. It is demonstrated that the optimum relation between pressure ratio and turbine entry temperature depends strongly on whether the optimization is carried out at fixed pressure ratios or at fixed temperatures. All previous references seem to have considered only the latter method.

Assessment of Real Gas Effects on Approximate and Boundary Layer Equations for Hypersonic Laminar Flow over Axisymmetric Bodies

2014 ◽  
Vol 1016 ◽  
pp. 534-539
Author(s):  
Ramin Kamali Moghadam ◽  
Seyed Amir Hosseini
Keyword(s):  
Boundary Layer ◽  
Body Region ◽  
Integral Matrix ◽  
Real Gas ◽  
Boundary Layer Equations ◽  
Blunt Bodies ◽  
Real Gas Effects ◽  
Efficient Code ◽  
Computational Procedures ◽  
Axisymmetric Bodies

Two efficient computational procedures based on the boundary layer equations and approximate relations areassessedin prediction of the laminar hypersonic flowfield for both the perfect gas and equilibrium air around the axisymmetric blunt body configurations. For the boundary layer procedure, the boundary layer equationsutilize the integral matrix solution algorithm for the blunt nose and after body region by using a space marching technique. The integral matrix procedure able us to create accurate and smooth results using the minimum grid in the boundary layer and minimize the computational costs. Applying the approximate method creates a robust and efficient code for heating calculations over the blunt bodies which flies in hypersonic regimes. These algorithms are highly appropriate to design of hypersonic reentry vehicles. The effects of real gas on the flowfield characteristics are also studied in two procedures.

An Assessment of Shock-Disturbances Interaction Considering Real Gas Effects

2018 ◽  
Vol 141 (1) ◽  
Author(s):  
K. Hejranfar ◽  
S. Rahmani
Keyword(s):  
Theoretical Analysis ◽  
The Body ◽  
Mathematical Function ◽  
Perfect Gas ◽  
Real Gas ◽  
One Dimensional ◽  
Wide Range ◽  
Different Types ◽  
Real Gas Effects ◽  
Analytical Relations

In this study, a theoretical analysis is performed to assess the interaction of freestream disturbances with a plane normal shock considering real gas effects. Such effects are important in a field with high velocities and high temperatures. To perform the theoretical analysis, the downstream disturbances field is expressed as a mathematical function of the upstream one by incorporating real gas effects in the formulation. Here, the linearized one-dimensional perturbed unsteady Euler equations are used for the classification of the downstream/upstream disturbances field and the linearized one-dimensional perturbed Rankine–Hugoniot equations are applied to provide a relationship between the disturbances field of two sides of the shock. To incorporate real gas effects in the formulation, real gas relations and equilibrium air curve-fits are used in the resulting system of equations. The general formulation presented here may be simplified to derive Morkovin's formulation by the perfect gas assumption. The magnitudes of downstream disturbances field resulting from different types of upstream disturbances field (entropy wave and fast/slow acoustic waves) with the shock are expressed by appropriate analytical relations. Results for different disturbance variables are presented for a wide range of upstream Mach number considering real gas effects and compared with those of the perfect gas and some conclusions are made. The effects of the presence of body are also studied theoretically and the analytical relations for the magnitude of the pressure disturbance at the body for different types of upstream disturbances field considering real gas effects are provided and their results are presented and discussed.

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