scholarly journals Investigation of gasdynamic parameters of a supersonic flow near a body at various Mach numbers

2020 ◽  
Vol 1697 ◽  
pp. 012240
Author(s):  
A V Shevchenko ◽  
A S Yuriev ◽  
S A Poniaev ◽  
T A Zhitnikov ◽  
E B Panfilov ◽  
...  
Keyword(s):  
Supersonic Flow ◽  
Mach Numbers

Computational study of unsteady gas flow in the combustion chamber of a ramjet engine with heat transfer

2021 ◽  
pp. 50-61
Author(s):  
Надежда Петровна Скибина
Keyword(s):  
Heat Transfer ◽  
Numerical Simulation ◽  
Supersonic Flow ◽  
Combustion Chamber ◽  
Gas Flow ◽  
Stokes Equations ◽  
Heat Conducting ◽  
Measuring Device ◽  
Ramjet Engine ◽  
Mach Numbers

Проведено численное исследование нестационарного турбулентного сверхзвукового течения в камере сгорания прямоточного воздушно-реактивного двигателя. Описана методика экспериментального измерения температуры на стенке осесимметричного канала в камере сгорания двигателя. Математическое моделирование обтекания исследуемой модели двигателя проводилось для скоростей набегающего потока M = 5 ... 7. Начальные и граничные условия задачи соответствовали реальному аэродинамическому эксперименту. Проанализированы результаты численного расчета. Рассмотрено изменение распределения температуры вдоль стенки канала с течением времени. Проведена оценка согласованности полученных экспериментальных данных с результатами математического моделирования. Purpose. The aim of this study is a numerical simulation of unsteady supersonic gas flow in a working path of ramjet engine under conditions identical to aerodynamic tests. Free stream velocity corresponding to Mach numbers M=5 ... 7 are considered. Methodology. Presented study addresses the methods of physical and numerical simulation. The probing device for thermometric that allows to recording the temperature values along the wall of internal duct was proposed. To describe the motion of a viscous heat-conducting gas the unsteady Reynolds averaged Navier - Stokes equations are considered. The flow turbulence is accounted by the modified SST model. The problem was solved in ANSYS Fluent using finite-volume method. The initial and boundary conditions for unsteady calculation are set according to conditions of real aerodynamic tests. The coupled heat transfer for supersonic flow and elements of ramjet engine model are realized by setting of thermophysical properties of materials. The reliability testing of numerical simulation has been made to compare the results of calculations and the data of thermometric experimental tests. Findings. Numerical simulation of aerodynamic tests for ramjet engine was carried out. The agreement between the results of numerical calculations and experimental measurements for the velocity in the channel under consideration was obtained; the error was shown to be 2%. The temperature values were obtained in the area of contact of the supersonic flow with the surface of the measuring device for the external incident flow velocities for Mach numbers M = 5 ... 7. The process of heating the material in the channel that simulated the section of the engine combustion chamber was analyzed. The temperature distribution was studied depending on the position of the material layer under consideration relative to the contact zone with the flow. Value. In the course of the work, the fields of flow around the model of a ramjet engine were obtained, including the region of supersonic flow in the inner part of axisymmetric channel. The analysis of the temperature fields showed that to improve the quality of the results, it is necessary to take into account the depth of the calorimetric sensor. The obtained results will be used to estimate the time of interaction of the supersonic flow with the fuel surface required to reach the combustion temperature.

RESEARCH OF CONICAL BODIES INTERACTIONS WITH IRREGULAR SUPERSONIC FLOW AT SMALL ATTACK ANGLES

2020 ◽  
Author(s):  
Евгений Алексеевич Прокопенко ◽  
Артем Васильевич Шевченко ◽  
Сергей Алексеевич Яшков ◽  
Игорь Анатольевич Дема ◽  
Тимофей Андреевич Житников ◽  
...  
Keyword(s):  
Supersonic Flow ◽  
Wind Tunnel ◽  
Stokes Equations ◽  
Navier Stokes ◽  
Perfect Gas ◽  
Navier Stokes Equations ◽  
Mach Numbers ◽  
Experimental Researches

В статье представлены теоретические и экспериментальные исследованиях сверхзвукового потока вблизи конических тел при различных числах Маха. Экспериментальные исследования выполнены с помощью сверхзвуковой атмосферно-вакуумной аэродинамической трубы Военно-космической академии имени А.Ф.Можайского. В основу теоретического исследования положена модель вязкого совершенного газа, описываемая уравнениями Навье-Стокса. The article presents theoretical and experimental researches of supersonic flow near conical bodies at various Mach numbers. Experimental researches were carried out using a supersonic atmospheric-vacuum wind tunnel of Mozhaisky Military Space Academy. The theoretical researches is based on the model of a viscous perfect gas described by the Navier-Stokes equations.

Transonic and Low Supersonic Flow Losses of Two Steam Turbine Blades at Large Incidences

2004 ◽  
Vol 126 (6) ◽  
pp. 966-975 ◽  
Author(s):  
S.-M. Li ◽  
T.-L. Chu ◽  
Y.-S. Yoo ◽  
W. F. Ng
Keyword(s):  
Experimental Data ◽  
Supersonic Flow ◽  
Turbine Blades ◽  
Supersonic Flows ◽  
Relative Increase ◽  
Pressure Probe ◽  
Flow Losses ◽  
New Correlation ◽  
Turbine Cascades ◽  
Mach Numbers

A linear cascade experiment was conducted to investigate transonic and low supersonic flow losses of two nozzle blades for the steam turbines. In the experiment, flow incidences were changed from −34° to 35° and exit Mach numbers were varied from 0.60 to 1.15. Tests were conducted at Reynolds numbers between 7.4×105 and 1.6×106. Flow visualization techniques, such as shadowgraph, Schlieren, and surface color oil were used to document the flows. Measurements were made by using downstream traverses with Pitot probe, upstream total pressure probe, and sidewall static pressure taps. The losses were found to be rather constant at subsonic flows. At transonic and low supersonic flows, the losses increased steeply. The maximum relative increase of the losses was near 700% when the Mach numbers increased from 0.6 to 1.15. However, the maximum relative increase of the losses was only about 100% due to very large variation of incidences. It is important to note that the effect of Mach numbers on losses was much greater than that due to the very large incidences for the transonic and low supersonic flows. A frequently used loss correlation in the literature is found not suitable to predict the losses of the tested blades for the transonic and low supersonic flows. From the current experimental data and some data in the literature, a new correlation for the shock related losses is proposed for transonic and low supersonic flows of turbine cascades. Comparison is made among the existing correlation and the new correlation, as well as the data of the current two cascades and other three turbine cascades in the literature. Improved agreement with the experimental data of the five cascades is obtained by using the new correlation as compared with the prediction by using the frequently used loss correlation in the literature.

Base Pressure Measurements on Elliptic Cones in Supersonic Flow

1967 ◽  
Vol 18 (3) ◽  
pp. 298-307 ◽  
Author(s):  
W. Stahl ◽  
H. Grauer-Carstensen
Keyword(s):  
Boundary Layer ◽  
Supersonic Flow ◽  
Reasonable Agreement ◽  
Flow Theory ◽  
Reynolds Numbers ◽  
Base Flow ◽  
Base Pressure ◽  
Pressure Measurements ◽  
Pressure Coefficients ◽  
Mach Numbers

SummaryAt the Aerodynamische Versuchsanstalt Gottingen (AVA), base pressure measurements were made on five elliptic cones. The ratios of the axes of the ellipses were: 1:12, 1:3, 1:1, 3:1, and 12:1. All the cones had the same volume and the same length. The investigations were carried out for Mach numbers M∞=1·50, 1·73, and 1·98 at angles of incidence between about —2 degrees and about 8 degrees. Reynolds numbers, based on a mean length, lm, varied from 2·5×106 to 3·0×106; the boundary layer approaching the base was turbulent. The base pressure coefficients are given as a function of geometry. Some of the results were compared with the base-flow theory of Korst and reasonable agreement was found.

Measurements of Friction in a Pipe for Subsonic and Supersonic Flow of Air

1946 ◽  
Vol 13 (2) ◽  
pp. A91-A100 ◽  
Author(s):  
J. H. Keenan ◽  
E. P. Neumann
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Supersonic Flow ◽  
Coefficient Of Friction ◽  
Reynolds Numbers ◽  
The Coefficient Of Friction ◽  
Mach Numbers ◽  
Pipe Inlet ◽  
Oblique Shocks

Abstract The apparent friction coefficient was determined experimentally for the flow of air through smooth pipes at subsonic and supersonic velocities. Values of the Mach number ranged from 0.27 to 3.87, and of Reynolds number from 1 × 105 to 8.7 × 105. In supersonic flow the results were found to be strongly influenced by the presence of oblique shocks formed at the junction of nozzle and pipe. The effect of these shocks on the coefficient of friction was determined. Nozzle forms were devised which eliminated the shocks and their effects. It was found that at distances from the pipe inlet greater than 50 diam the apparent coefficient of friction for compressible flow at Mach numbers greater or less than 1 is approximately equal, for equal Reynolds numbers, to the coefficient of friction for incompressible flow with completely developed boundary layer. Mach numbers greater than 1 are rarely maintained for lengths greater than 50 diam. For attainable lengths, the coefficient of friction is a function of the ratio of length to diameter and the Reynolds number, with the Mach number at entrance determining the maximum attainable length.

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