Wall Pressure Measurements in a Convergent–Divergent Nozzle with Varying Inlet Asymmetry

2016 ◽  
Vol 33 (2) ◽  
Author(s):  
C. Senthilkumar ◽  
S. Elangovan ◽  
E. Rathakrishnan
Keyword(s):  
Mach Number ◽  
Supersonic Flow ◽  
Flow Separation ◽  
Static Pressure ◽  
Flow Characteristics ◽  
Pressure Measurements ◽  
Separation Region ◽  
Number Ratio ◽  
Inlet Angle ◽  
Nozzle Inlet

AbstractIn this paper, flow separation of a convergent–divergent (C-D) nozzle is placed downstream of a supersonic flow delivered from Mach 2.0 nozzle is investigated. Static pressure measurements are conducted using pressure taps. The flow characteristics of straight and slanted entry C-D nozzle are investigated for various NPR of Mach 2.0 nozzle. The effect of asymmetry at inlet by providing 15°, 30°, 45° and 57° cut is analyzed. Particular attention is given to the location of the shock within the divergent section of the test nozzle. This location is examined as a function both NPR of Mach 2.0 nozzle and test nozzle inlet angle. Some of the measurements are favorably compared to previously developed theory. A Mach number ratio of 0.81 across the flow separation region was obtained.

Method for Providing Warning of the Onset of Buffeting, Stalling and Other Undesirable Effects of Flow Separation

1958 ◽  
Vol 62 (573) ◽  
pp. 674-676 ◽  
Author(s):  
D. W Holder ◽  
H. H. Pearcey
Keyword(s):  
Shock Wave ◽  
Recent Work ◽  
Flow Separation ◽  
Static Pressure ◽  
Separated Flow ◽  
Leading Edge ◽  
Alternative Methods ◽  
Pressure Measurements ◽  
Aircraft Wing ◽  
Trailing Edge

A Method is described for providing warning of the onset of undesirable effects produced by flow separation on an aircraft wing. It is based on static pressure measurements near the trailing edge and appears to have advantages over alternative methods.Recent work has suggested that certain effects of flow separation such as buffeting, aileron buzz and undesirable changes in loading, occur when a “ bubble ” of separated flow originating at the leading edge of a wing, or at a shock wave on its surface, first becomes sufficiently large to affect the flow at the trailing edge.

A Method for Correcting Wall Pressure Measurements in Subsonic Compressible Flow

1991 ◽  
Vol 113 (2) ◽  
pp. 256-260 ◽  
Author(s):  
C. Ducruet
Keyword(s):  
Boundary Layer ◽  
Experimental Investigation ◽  
Mach Number ◽  
Compressible Flow ◽  
Velocity Gradient ◽  
Static Pressure ◽  
Subsonic Flow ◽  
Diameter Ratio ◽  
Wall Pressure ◽  
Pressure Measurements

A theoretical and experimental investigation has been made of the static pressure hole problem in subsonic flow. Thanks to a linearization, the effects of the boundary layer, of the velocity gradient and of the wall curvature could be separated so that a formula of correction containing three influence functions has been obtained. These functions were determined in the case of practical requirements by means of experiments made on appropriate models for two values of the depth-to-diameter ratio and for at least three values of the Mach number. Then, the method of correction has been applied to the flow around an airfoil at zero angle of attack.

Experimental Method for Combustion Efficiency Calculation in a Reheat Duct

1988 ◽  
Vol 110 (4) ◽  
pp. 690-694 ◽  
Author(s):  
A. Cadiou
Keyword(s):  
Cross Section ◽  
Static Pressure ◽  
Flow Characteristics ◽  
Combustion Efficiency ◽  
Pressure Measurements ◽  
Two Dimensional ◽  
Combustion Chambers ◽  
Outlet Cross Section ◽  
Gas Sampling ◽  
Flame Holder

An efficient method for the experimental measurement of the combustion efficiency in a reheat duct has been developed at ONERA. Such a method is useful because numerous reheat tests are necessary to study the effect of geometry and flow characteristics on reheat performances. Static pressure measurements along the duct and gas sampling in its outlet cross section are the basis of this downstream-to-upstream method. Experimental results with a tri-annular V-gutter flame holder are presented. These results are also used for comparison with theoretical two-dimensional calculations applied to reheat ducts that ultimately may reduce the number of experiments necessary for the development of reheat combustion chambers.

scholarly journals A preliminary study on static pressure probe mobile measurement method for flow field in transonic wind tunnel

2021 ◽  
Vol 39 (4) ◽  
pp. 786-793
Author(s):  
Haijun Deng ◽  
Bo Xiong ◽  
Xinfu Luo ◽  
Shaozun Hong ◽  
Qi Liu ◽  
...  
Keyword(s):  
Mach Number ◽  
Wind Tunnel ◽  
Flow Field ◽  
Static Pressure ◽  
Wind Tunnel Test ◽  
Flow Characteristics ◽  
Pressure Probe ◽  
Number Distribution ◽  
Mach Number Distribution ◽  
Transonic Wind Tunnel

The axial Mach number distribution of the core flow for model in a transonic wind tunnel is an important index to evaluate the performance of the flow field, which is usually measured by the centerline probe. In order to simulate the incoming flow characteristics without interference, the probe will extend from the support section to the shrinkage section, so the probe usually must has longer inches, more static pressure measuring points and smaller blockage requirements. In order to study the influence of the points of the centerline probe on the uniformity distribution of flow field, a new static pressure probe is designed, which is smaller and shorter than the centerline probe. On the basis of the stability of the flow field, the Mach number distribution of the flow field measured by the static pressure probe which is driven by the moving measuring mechanism. The characteristics of the measured values are studied by wind tunnel test. The results show that: when Ma ≤ 0.95, the overall distribution and value of Mach number obtained by the static pressure probe is basically the same as those obtained by the centerline probe, but some flow field details, which mainly shows that Mach number of the static pressure probe has smaller fluctuation, higher accuracy and better uniformity index.

Investigation of Flow Separation Characteristics in a Pump as Turbines Impeller Under the Best Efficiency Point Condition

2021 ◽  
Vol 143 (6) ◽  
Author(s):  
Tong Lin ◽  
Xiaojun Li ◽  
Zuchao Zhu ◽  
Renhua Xie ◽  
Yanpi Lin
Keyword(s):  
Flow Separation ◽  
Topological Structure ◽  
Static Pressure ◽  
Flow Characteristics ◽  
Internal Flow ◽  
Surface Friction ◽  
Suction Side ◽  
Peak Position ◽  
Transient Pressure ◽  
Point Condition

Abstract The impeller, which is the main energy conversion component of a pump as turbine (PAT), is designed for pumping mode, and its internal flow characteristics are quite complicated even at the best efficiency point (BEP) of the turbine mode. This study aims to investigate the flow separation characteristics in a PAT impeller under the BEP condition by numerical method. The hydraulic performance and transient pressure characteristics of PAT predicted numerically were verified through experimentation. The surface friction lines and flow topological structure were applied to diagnose the flow separation at the surface of the blade. The relationship between flow topological structure and vortex in the impeller and static pressure at the blade were analyzed. Analysis results show that the backflow and open flow separation are observed significantly in the leading region and near the shroud of the trailing region of suction side. The passage vortex always appears near the spiral node. The saddle point and spiral node correspond to the peak position of adverse pressure and the lowest position between two peak values of the static pressure of the blade, respectively. The inflow conditions of blade and shape of the trailing edge significantly influence the flow separations in the impeller.

scholarly journals Возникновение усилий на клине при взаимодействии скользящего разряда со сверхзвуковым потоком

2021 ◽  
Vol 47 (9) ◽  
pp. 33
Author(s):  
В.М. Бочарников ◽  
В.В. Голуб ◽  
А.С. Савельев
Keyword(s):  
Shock Wave ◽  
Reynolds Number ◽  
Mach Number ◽  
Supersonic Flow ◽  
Wind Tunnel ◽  
Static Pressure ◽  
Discharge Energy ◽  
Pulse Heat ◽  
The Impact ◽  
Sliding Discharge

It was investigated generation of forces on the wedge during the interaction of a sliding discharge in the supersonic flow. Sliding discharge was generated in a supersonic atmospheric-vacuum wind tunnel with a flow Mach number M = 2, Reynolds number of about 106, and static pressure in the flow p = 0.15 bar. It was found that the impact of the discharge is not limited by the appearance of a shock wave, and also significantly changes the properties of the flow around the wedge due to pulse heat generation. As the discharge energy increases, the force from its impact on the wedge also increases. Force produced by the sliding discharge in the supersonic flow is several times greater than in the quiescent air regardless of its pressure: 0.15 bar or 1 bar.

Mathematical and Physical Analysis of the Effect of Conical and Detached Shock Waves at the Tip of a Static Probe in an Experimental Chamber

2021 ◽  
Vol 105 (1) ◽  
pp. 627-635
Author(s):  
Pavla Šabacká ◽  
Jiří Maxa ◽  
Anna Maxová
Keyword(s):  
Supersonic Flow ◽  
Static Pressure ◽  
Experimental Chamber ◽  
Physical Analysis ◽  
Pressure Measurements ◽  
Scientific Instruments ◽  
Electronic Technology ◽  
Vacuum Chambers ◽  
University Of Technology ◽  
Flow Through

As part of the research in the field of pumping vacuum chambers in the Environmental Electron Microscope, research on supersonic flow through apertures is being carried out at the Department of Electrical and Electronic Technology of the Brno University of Technology in cooperation with the Institute of Scientific Instruments of the CAS. This paper deals with the influence of the shape of the static probe cone design for static pressure measurements in the supersonic flow regime in the Experimental Chamber. The cone of the probe has an effect on the shape of the shock wave, which significantly influences the detected static pressure value.

Static-pressure probes that are theoretically insensitive to pitch, yaw and Mach number

1970 ◽  
Vol 44 (3) ◽  
pp. 513-528 ◽  
Author(s):  
A. M. O. Smith ◽  
A. B. Bauer
Keyword(s):  
Mach Number ◽  
Cross Sections ◽  
Flow Separation ◽  
Potential Flow ◽  
Static Pressure ◽  
Angle Of Attack ◽  
Test Results ◽  
Cross Sectional

The idea of distributing static probe cross-sectional areas so as to render the probe insensitive to Mach number is combined here with that of using non-circular cross-sections to render probes insensitive to yaw and angle of attack. Appropriate non-circular cross-sections are described in detail, and a general means of designing blunt or slender probes to have zero sensitivity to yaw and angle of attack in potential flow is described. Four experimental probes have been tested, and test results are presented. These results show that the probes are quite insensitive to yaw and angle of attack within certain limiting angles, which are assumed to correspond to the onset of flow separation.

Improvement of the Performance of a Supersonic Nozzle by Riblets

2000 ◽  
Vol 122 (3) ◽  
pp. 585-591 ◽  
Author(s):  
Kazumi Tsunoda ◽  
Tomohiko Suzuki ◽  
Toshiaki Asai
Keyword(s):  
Mach Number ◽  
Pressure Loss ◽  
Static Pressure ◽  
Supersonic Nozzle ◽  
Pressure Rise ◽  
Internal Flow ◽  
Stagnation Pressure ◽  
Separation Region ◽  
Loss Reduction ◽  
Riblet Surface

This paper describes an experimental study of supersonic internal flow over a riblet surface mounted on a channel wall to reduce pressure loss and improve the performance of a supersonic nozzle. The magnitude of the static pressure in the pressure-rise region observed in channels with riblet surface became lower than that for a smooth surface, and the significance of its difference was indicated by uncertainty analysis estimated at 95 percent coverage. The Mach number distributions obtained by traversing a Pitot-tube showed that the separation point moved downstream and the size of the separation region became small when using riblets. Furthermore, it was found that the stagnation pressure loss reduction was as large as 56 percent in the uniform supersonic flow field at a Mach number of 2.0, and 29 percent in the separation region. [S0098-2202(00)00103-6]

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