Effect of orifice spacing on twin circular parallel compressible jets

2021 ◽  
Vol 0 (0) ◽  
Author(s):  
K. B. V. Satya Prakash ◽  
P. Lovaraju ◽  
E. Rathakrishnan
Keyword(s):  
Mach Number ◽  
Orifice Diameter ◽  
Jet Mixing ◽  
Circular Jets ◽  
Single Jet ◽  
Mach Numbers ◽  
Exit Mach Number ◽  
Center Distance

Abstract The interaction of Mach 0.5, 0.8, and 1.0, parallel, twin circular jets issuing from orifices with center-to-center spacing S/D, where S is the center-to-center distance and D is orifice diameter, 2, 4 and 6 has been investigated experimentally. The characteristics of twinjets are analyzed based on the centerline Mach number decay, exit Mach number and ratio of orifice spacing. As the spacing between the orifice increases, the maximum Mach number point of the combined jet moves downstream. For the Mach numbers studied it is found that as the S/D increases the effect of the counter-rotating vortices on jet mixing decreases. The rate of the twinjet interaction also decreases with S/D increase. As the jets propagate downstream their center-to-center distance decreases continuously and the jets merge to become single jet, for all S/D studied.

Compressible Flow Ejectors: Part II—Flow Field Measurements and Analysis

1974 ◽  
Vol 96 (3) ◽  
pp. 282-288 ◽  
Author(s):  
K. R. Hedges ◽  
P. G. Hill
Keyword(s):  
Experimental Study ◽  
Mach Number ◽  
Mass Flow ◽  
Field Measurements ◽  
Flow Ratio ◽  
Flow Rates ◽  
Jet Mixing ◽  
Mass Flow Rates ◽  
Exit Mach Number ◽  
Mass Flow Ratio

An experimental study has been made of compressible jet mixing in an axisymmetric ejector of converging-diverging geometry. The mass flow ratio was in the range 1.3 to 2.6 and the nozzle exit Mach number was 1.82. Ejector performance characteristics were obtained as well as measurements of pressure and velocity distribution over a range of mass flow rates. The experimental results were used to test the reliability of the analytical model of the flow described in Part I of the paper.

Effect of Turbine Airfoil Shape on Aerodynamic Losses for Turbine Airfoils Operating Under Transonic Conditions

2011 ◽  
Author(s):  
Santosh Abraham ◽  
Kapil Panchal ◽  
Srinath V. Ekkad ◽  
Wing Ng ◽  
Barry J. Brown ◽  
...  
Keyword(s):  
Mach Number ◽  
Flow Field ◽  
Pressure Loss ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Airfoil Surface ◽  
Mach Numbers ◽  
Exit Mach Number ◽  
Turbine Airfoils ◽  
Turbine Design

Profile and secondary loss correlations have been developed and improved over the years to include the induced incidence and leading edge geometry and to reflect recent trends in turbine design. All of these investigations have resulted in better understanding of the flow field in turbine passages. However, there is still insufficient data on the performance of turbine airfoils with high turning angles operating at varying incidence angles at transonic Mach numbers. The paper presents detailed aerodynamic measurements for three different turbine airfoils with similar turning angles but different aerodynamic shapes. Midspan total pressure loss, secondary flow field, and static pressure measurements on the airfoil surface in the cascades are presented and compared for the three different airfoil sets. The airfoils are designed for the same velocity triangles (inlet/exit gas angles and Mach number). Airfoil curvature and true chord are varied to change the loading vs. chord. The objective is to investigate the type of loading distribution and its effect on aerodynamic performance (pressure loss). Measurements are made at +10, 0 and −10 degree incidence angles for high turning turbine airfoils with ∼127 degree turning. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. In order to attain a ratio of inlet Mach number to exit Mach number that is representative to that encountered in a real engine, the exit span is increased relative to the inlet span. This results in one end wall diverging from inlet to exit at a 13 degree angle, which simulates the required leading edge loading as seen in an engine. 3D viscous compressible CFD analysis was carried out in order to compare the results with experimentally obtained values and to further investigate the flow characteristics of the airfoils under study.

Numerical investigation of critical lip thickness of subsonic co-flowing jet

2021 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Naren Shankar R. ◽  
Ganesan V.G. ◽  
Dilip Raja N. ◽  
Sathish Kumar K. ◽  
Vijayaraja K.
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Static Pressure ◽  
Velocity Ratio ◽  
Critical Limit ◽  
Number Range ◽  
Content Type ◽  
Jet Mixing ◽  
Pressure Decay ◽  
Mach Numbers

Purpose The effect of increasing lip thickness (LT) and Mach number on subsonic co-flowing Jet (CFJ) decay at subsonic and correctly expanded sonic Mach numbers has been analysed experimentally and numerically in this study. This study aims to a critical LT below which mixing enhances and above which mixing inhibits. Design/methodology/approach LT is the distance, separating the primary nozzle and the secondary duct, present in the co-flowing nozzle. The CFJ with LT ranging from 2 mm to 150 mm at jet exit Mach numbers of 0.6, 0.8 and 1.0 were studied in detail. The CFJ with 2 mm LT is used for comparison. Centreline total pressure decay, centreline static pressure decay and near field flow behaviour were analysed. Findings The result shows that the mixing enhances until a critical limit and a further increase in the LT does not show any variation in the jet mixing. Beyond this critical limit, the secondary jet has a detrimental effect on the primary jet, which deteriorates the process of mixing. The CFJ within the critical limit experiences a significantly higher mixing. The effect of the increase in the Mach number has marginal variation in the total pressure and significant variation in static pressure along the jet axis. Practical implications In this study, the velocity ratio (VR) is maintained constant and the bypass ratio (BR) was varied from low value to very high values for subsonic and correctly expanded sonic. Presently, commercial aircraft engine operates under these Mach numbers and low to ultra-high BR. Hence, the present study becomes essential. Originality/value This is the first effort to find the critical value of LT for a constant VR for a Mach number range of 0.6 to 1.0, compressible CFJ. The CFJs with constant VR of unity and varying LT, in these Mach number range, have not been studied in the past.

Tip Leakage Losses in Subsonic and Transonic Blade-Rows

2011 ◽  
Author(s):  
Andrew P. S. Wheeler ◽  
Theodosios Korakianitis ◽  
Shashimal Banneheke
Keyword(s):  
Mach Number ◽  
Three Dimensional ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Blade Loading ◽  
Tip Leakage ◽  
Leakage Flows ◽  
Mach Numbers ◽  
Exit Mach Number ◽  
Tip Leakage Flows

In this paper the effect of blade-exit Mach number on unshrouded turbine tip-leakage flows is investigated. Previously published experimental data of a high-pressure turbine blade are used to validate a CFD code, which is then used to study the tip-leakage flow at blade-exit Mach numbers from 0.6 to 1.4. Three-dimensional calculations are performed of a flat-tip and a cavity-tip blade. Two-dimensional calculations are also performed to show the effect of various squealer-tip geometries on an idealized tip-flow. The results show that as the blade-exit Mach number is increased the tip leakage flow becomes choked. Therefore the tip-leakage flow becomes independent of the pressure difference across the tip and hence the blade-loading. Thus the effect of the tip-leakage flow on overall blade loss reduces at blade-exit Mach numbers greater than 1.0. The results suggest that for transonic blade-rows it should be possible to raise blade loading within the tip region without increasing tip-leakage loss.

Influence of Loading Distribution on the Off-Design Performance of High-Pressure Turbine Blades

2006 ◽  
Author(s):  
D. Corriveau ◽  
S. A. Sjolander
Keyword(s):  
High Pressure ◽  
Mach Number ◽  
Turbine Blades ◽  
Operating Conditions ◽  
Design Flow ◽  
Superior Performance ◽  
High Pressure Turbine ◽  
Design Performance ◽  
Mach Numbers ◽  
Exit Mach Number

Linear cascade measurements for the aerodynamic performance of a family of three transonic, high-pressure (HP) turbine blades have been presented previously by the authors. The airfoils were designed for the same inlet and outlet velocity triangles but varied in their loading distributions. The previous papers presented results for the design incidence at various exit Mach numbers, and for off-design incidence at the design exit Mach number of 1.05. Results from the earlier studies indicated that by shifting the loading towards the rear of the airfoil an improvement in the profile loss performance of the order of 20% could be obtained near the design Mach number at design incidence. Measurements performed at off-design incidence, but still at the design Mach number, showed that the superior performance of the aft-loaded blade extended over a range of incidence from about −5.0° to +5.0° relative to the design value. For the current study, additional measurements were performed at off-design Mach numbers from about 0.5 to 1.3 and for incidence values of −10.0°, +5.0° and + 10.0° relative to design. The corresponding Reynolds numbers, based on outlet velocity and true chord, varied from roughly 4 × 105 to 10 × 105. The measurements included midspan losses, blade loading distributions and base pressures. In addition, two-dimensional Navier-Stokes computations of the flow were performed to help in the interpretation of the experimental results. The results show that the superior loss performance of the aft-loaded profile, observed at design Mach number and low values of off-design incidence, does not extend readily to off-design Mach numbers and larger values of incidence. In fact, the measured midspan loss performance for the aft-loaded blade was found to be inferior to, or at best equal to, that of the baseline, mid-loaded airfoil at most combinations of off-design Mach number and incidence. However, based on the observations made at design and off-design flow conditions, it appears that aft-loading can be a viable design philosophy to employ in order to reduce the losses within a blade row provided the rearward deceleration is carefully limited. The loss performance of the front-loaded blade is inferior or at best equal to that of the other two blades for all operating conditions. As such, it appears that there is no advantage to front loading the airfoil for transonic high-pressure turbine blades. The results also provide a significant addition to the data available in the open literature on the off-design performance of transonic HP turbine airfoils.

Effect of Airfoil Shape and Turning Angle on Turbine Airfoil Aerodynamic Performance at Transonic Conditions

2011 ◽  
Author(s):  
Santosh Abraham ◽  
Kapil Panchal ◽  
Srinath V. Ekkad ◽  
Wing Ng ◽  
Barry J. Brown ◽  
...  
Keyword(s):  
Mach Number ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Flow Characteristics ◽  
Airfoil Surface ◽  
Turning Angle ◽  
Gas Turbine Blades ◽  
Mach Numbers ◽  
Exit Mach Number ◽  
Turbine Airfoils

Performance data for high turning gas turbine blades under transonic Mach numbers is significantly lacking in literature. Performance of three gas turbine airfoils with varying turning angles at transonic flow conditions was investigated in this study. Midspan total pressure loss, secondary flow field and static pressure measurements on the airfoil surface in a linear cascade setting were measured. Airfoil curvature and true chord were varied to change the loading vs. chord for each airfoil. Airfoils A, D and E are designed to operate at different velocity triangles. Velocity triangle requirements (inlet/exit Mach number and gas angles) come from 1D and 2D models that include calibrated loss systems. One of the goals of this study was to use the experimental data to confirm/refine loss predictions for the effect of various Mach numbers and gas turning angles. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. The airfoil turning angle ranges from 120° to 138°. A realistic inlet/exit Mach number ratio, that is representative of that seen in a real engine, was obtained by reducing the inlet span with respect to the exit span of the airfoil, thereby creating a quasi 2D cascade. In order to compare the experimental results and study the detailed flow characteristics, 3D viscous compressible CFD analysis was also carried out.

Investigation of Effect of End Wall Contouring Methods on a Transonic Turbine Blade Passage

2011 ◽  
Author(s):  
Kapil Panchal ◽  
Santosh Abraham ◽  
Srinath V. Ekkad ◽  
Wing Ng ◽  
Barry J. Brown ◽  
...  
Keyword(s):  
Mach Number ◽  
Turbine Blade ◽  
Surface Topology ◽  
Design Parameters ◽  
Loss Reduction ◽  
Blade Passage ◽  
Transonic Turbine ◽  
Mach Numbers ◽  
Exit Mach Number ◽  
End Wall

End wall contouring has been widely studied during past two decades for secondary loss reduction in turbine passages. Recent non-axisymmetric end wall contouring methods have shown more promise for loss reduction as compared to the axisymmetric end wall contouring methods used in initial studies. The end wall contouring methods have shown definite promise, especially, for the turbine passages at low design exit Mach numbers. A class of methods exists in the literature where the end wall surface is defined by using a combination of two curves. These curves specify surface topology variation in streamwise and pitchwise directions. Another class of methods depends on surface contour optimization, in which the modification of surface contours is achieved by changing the control point locations that define the surface topology. A definitive, passage design parameter based method of contouring is still not available. However, a general guideline for the trend of contour variation, along pitchwise and streamwise direction, can certainly be extrapolated from the existing literature. It is not clear, however, whether such a trend can be fitted to any blade profile to achieve, least of all a nonoptimum but a definite, reduction in losses. Moreover, almost all of the existing studies have focused on end wall contouring of passages with low exit Mach numbers. Some researchers, indeed, have used blades designed for high turning and high exit Mach number. However, such studies were done at Mach number well below the intended design condition. A study of effect of end wall contouring on a high turning blade with high design exit Mach number is not available in open literature. The present study investigates the effect of application of three different types of end wall contouring methods through numerical simulation, on a high turning transonic turbine blade passage. The main contouring method is based on total loss reduction criterion which is described here in detail. The contouring methodology described here avoids the deficiency of current commercial mesh generation software in context of automated meshing and provides a robust end wall optimization methodology. The geometry that gives minimum SKE values is compared with this loss optimized geometry. Additionally, a normalized contoured surface topology was extracted from a previous study that has similar blade design parameters and this surface was fitted to the turbine passage under study in order to investigate the effect of such trend based surface fitting. This contour geometry has also been compared with the other two contour geometries. Aerodynamic response of these geometries has been compared in detail with the baseline case without any end wall contouring. A comparison of shape and location of end wall contours on aerodynamic performance has been provided. The results indicate that end wall contouring for transonic turbine blades may not result in as significant gains at design conditions as those claimed for low speed turbine passages in previous studies.

Design and Numerical Simulation of Convergent Divergent Nozzle

2016 ◽  
Vol 852 ◽  
pp. 617-624
Author(s):  
V. Ramji ◽  
Raju Mukesh ◽  
Inamul Hasan
Keyword(s):  
Mach Number ◽  
Computational Analysis ◽  
Turbulence Modeling ◽  
Pressure Ratio ◽  
Minimum Length ◽  
Ansys Fluent ◽  
Exit Pressure ◽  
Constant Height ◽  
Mach Numbers ◽  
Exit Mach Number

This works centers on the design of a De Laval (convergent - Divergent) nozzle to accelerate the flow to supersonic or hypersonic speeds and computational analysis of the same. An initial design of the nozzle is made from the method of characteristics. The coding was done in Matlab to obtain the contour of the divergent section for seven different exit Mach numbers viz. 2.5,3,3.5,4,4.5,5 and 5.5.To quantify variation in the minimum length of the nozzle divergent section with respect to the exit mach number, a throat of constant height (0.005m) and width (0.05m) was chosen for all the design. The area exit required for each mach no varying from 1 to 5.5 was plotted using isentropic relations and was also used to verify the exit area of the nozzle for each of those mach numbers. An estimate of the exit pressure ratio is obtained by using isentropic and normal shock relations. With this exit pressure ratio, a more refined verification is done by computational analysis using ANSYS Fluent software for a contour nozzle with exit Mach number 5.5. The spalart Allmaras and k-epsilon model were used for turbulence modeling.

Influence of Loading Distribution on the Off-Design Performance of High-Pressure Turbine Blades

2006 ◽  
Vol 129 (3) ◽  
pp. 563-571 ◽  
Author(s):  
D. Corriveau ◽  
S. A. Sjolander
Keyword(s):  
High Pressure ◽  
Mach Number ◽  
Turbine Blades ◽  
Reynolds Numbers ◽  
Operating Conditions ◽  
Design Flow ◽  
Superior Performance ◽  
Navier Stokes ◽  
Mach Numbers ◽  
Exit Mach Number

Linear cascade measurements for the aerodynamic performance of a family of three transonic, high-pressure (HP) turbine blades have been presented previously by the authors. The airfoils were designed for the same inlet and outlet velocity triangles but varied in their loading distributions. The previous papers presented results for the design incidence at various exit Mach numbers, and for off-design incidence at the design exit Mach number of 1.05. Results from the earlier studies indicated that by shifting the loading towards the rear of the airfoil an improvement in the profile loss performance of the order of 20% could be obtained near the design Mach number at design incidence. Measurements performed at off-design incidence, but still at the design Mach number, showed that the superior performance of the aft-loaded blade extended over a range of incidence from about −5.0deg to +5.0deg relative to the design value. For the current study, additional measurements were performed at off-design Mach numbers from about 0.5 to 1.3 and for incidence values of −10.0deg, +5.0deg, and +10.0deg relative to design. The corresponding Reynolds numbers, based on outlet velocity and true chord, varied from roughly 4×105 to 10×105. The measurements included midspan losses, blade loading distributions, and base pressures. In addition, two-dimensional Navier–Stokes computations of the flow were performed to help in the interpretation of the experimental results. The results show that the superior loss performance of the aft-loaded profile, observed at design Mach number and low values of off-design incidence, does not extend readily to off-design Mach numbers and larger values of incidence. In fact, the measured midspan loss performance for the aft-loaded blade was found to be inferior to, or at best equal to, that of the baseline, midloaded airfoil at most combinations of off-design Mach number and incidence. However, based on the observations made at design and off-design flow conditions, it appears that aft-loading can be a viable design philosophy to employ in order to reduce the losses within a blade row provided the rearward deceleration is carefully limited. The loss performance of the front-loaded blade is inferior or at best equal to that of the other two blades for all operating conditions.

scholarly journals Aerodynamics of a Transitioning Turbine Stator Over a Range of Reynolds Numbers

1998 ◽  
Author(s):  
R. J. Boyle ◽  
B. L. Lucci ◽  
V. G. Verhoff ◽  
W. P. Camperchioli ◽  
H. La
Keyword(s):  
Mach Number ◽  
Exhaust System ◽  
Reynolds Numbers ◽  
Inlet Pressure ◽  
Navier Stokes ◽  
Low Reynolds Numbers ◽  
Turbine Vane ◽  
Loss Coefficients ◽  
Mach Numbers ◽  
Exit Mach Number

Midspan aerodynamic measurements for a three vane-four passage linear turbine vane cascade are given. The vane axial chord was 4.45cm. Surface pressures and loss coefficients were measured at exit Mach numbers of 0.3, 0.7, and 0.9. Reynolds number was varied by a factor of six at the two highest Mach numbers, and by a factor often at the lowest Mach number. Measurements were made with and without a turbulence grid. Inlet turbulence intensities were less than 1% and greater than 10%. Length scales were also measured. Pressurized air fed the test section, and exited to a low pressure exhaust system. Maximum inlet pressure was two atmospheres. The minimum inlet pressure for an exit Mach number of 0.9 was one-third of an atmosphere, and at a Mach number of 0.3, the minimum pressure was half this value. The purpose of the test was to provide data for verification of turbine vane aerodynamic analyses, especially at low Reynolds numbers. Predictions obtained using a Navier-Stokes analysis with an algebraic turbulence model are also given.

Sign in / Sign up

Export Citation Format

Share Document