The Impingement of a Uniform, Axisymmetric, Supersonic Jet on a Perpendicular Flat Plate

1971 ◽  
Vol 22 (4) ◽  
pp. 403-420 ◽  
Author(s):  
J. H. Gummer ◽  
B. L. Hunt
Keyword(s):  
Mach Number ◽  
Flat Plate ◽  
Supersonic Jet ◽  
Theoretical Description ◽  
Alternative Methods ◽  
Number Range ◽  
Sonic Line ◽  
Pressure Distributions ◽  
Experimental Shock ◽  
Mach Numbers

SummaryThe impingement region produced by directing a uniform, axisymmetric, supersonic jet of air normally onto a large, flat plate has been investigated experimentally and theoretically for four jets in the Mach number range 1·64 to 2·77. A qualitative theoretical description of the flow in the neighbourhood of the sonic line is given. A single-strip version of the method of Polynomial Approximation and Integral Relations (PIR) is applied to the flow, using two alternative methods of determining the centre-line shock height. The PIR predictions are compared to experimental shock shapes and pressure distributions. It is found that a PIR method in which the sonic line is assumed to intersect the shock at the jet edge leads to very good agreement with experiment at the higher jet Mach numbers, but accuracy is much reduced at the lower Mach numbers, the shock height being in error by about 62 per cent at a jet Mach number of 1·64. A change in flow pattern at small nozzle-to-plate distances is reported.

The near wall jet of a normally impinging, uniform, axisymmetric, supersonic jet

1974 ◽  
Vol 66 (1) ◽  
pp. 159-176 ◽  
Author(s):  
J. C. Carling ◽  
B. L. Hunt
Keyword(s):  
Mach Number ◽  
Shock Layer ◽  
Supersonic Jet ◽  
Surface Flow ◽  
Wall Jet ◽  
Viscous Effects ◽  
Number Range ◽  
Axisymmetric Jet ◽  
Pressure Distributions ◽  
Near Wall

The near wall jet produced by directing a uniform axisymmetric jet of air normally onto a large flat plate has been investigated experimentally and theoretically for four jets in the Mach number range 1·64–2·77. Detailed measurements of the surface pressure and shadowgraph and surface flow pictures are presented. The results show that the mechanism which mainly determines the supersonic near wall jet is the jet-edge expansion and its reflexions from the sonic line and the wall-jet boundaries. The near wall jet is found to consist of an alternating series of expansion and recompression regions whose strengths depend on the jet Mach number and decay with distance. At Mach numbers of 2·4 and above, shock waves are observed in the first recompression region and at a Mach number of 2·77 the boundary layer separates locally. Further out, viscous effects become increasingly important and a constant-pressure shear flow is established at a distance which increases with jet Mach number. The application of the method of characteristics in an approximate manner reproduces a number of the features of the near wall jet which are observed experimentally.Pressure distributions obtained in the shock layer show that a stagnation bubble can occur and that its occurrence depends on factors such as the flow upstream of the nozzle. The wall-jet region is found to be largely independent of whether or not a bubble occurs in the shock layer.

Numerical Investigation of Impingement Heat Transfer on a Flat and Square Pin-Fin Roughened Plates

2013 ◽  
Author(s):  
Chaoyi Wan ◽  
Yu Rao ◽  
Xiang Zhang
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Flat Plate ◽  
Numerical Investigation ◽  
Pressure Loss ◽  
Transfer Characteristic ◽  
Number Range ◽  
Pin Fin ◽  
Pressure Distributions ◽  
Impingement Heat Transfer

A numerical investigation of the heat transfer characteristics within an array of impingement jets on a flat and square pin-fin roughened plate with spent air in one direction has been conducted. Four types of optimized pin-fin configurations and the flat plate have been investigated in the Reynolds number range of 15000–35000. All the computation results have been validated well with the data of published literature. The effects of variation of jet Reynolds number and different configurations on the distribution of the average and local Nusselt number and the related pressure loss have been obtained. The highest total heat transfer rate increased up to 162% with barely any extra pressure loss compared with that of the flat plate. Pressure distributions and streamlines have also been captured to explain the heat transfer characteristic.

Measured Pressure Distributions and Shock Shapes on a Simple Delta Wing

1981 ◽  
Vol 32 (3) ◽  
pp. 188-198 ◽  
Author(s):  
L.C. Squire
Keyword(s):  
Experimental Investigation ◽  
Mach Number ◽  
Delta Wing ◽  
Numerical Calculations ◽  
Test Case ◽  
Pressure Distributions ◽  
Shock System ◽  
Mach Numbers ◽  
Measured Pressure

SummaryThis note presents the results of an experimental investigation of the flow over a simple delta wing designed for a Mach number of 3.5. Complete pressure distributions were measured for incidences of 0°, 10° and 20° at Mach numbers of 2.5 and 3.5. A number of schlieren photographs of the shock system around the wing were obtained at the same conditions and surface streamline patterns were studied at M = 3.5. The measurements were made to support numerical calculations which use this wing as a test case.

Heat transfer from a hypersonic turbulent boundary layer on a flat plate

1973 ◽  
Vol 60 (2) ◽  
pp. 257-271 ◽  
Author(s):  
G. T. Coleman ◽  
C. Osborne ◽  
J. L. Stollery
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Reynolds Number ◽  
Friction Coefficient ◽  
Mach Number ◽  
Turbulent Boundary Layer ◽  
Flat Plate ◽  
Reasonable Agreement ◽  
Skin Friction Coefficient ◽  
Number Range

A hypersonic gun tunnel has been used to measure the heat transfer to a sharpedged flat plate inclined at various incidences to generate local Mach numbers from 3 to 9. The measurements have been compared with a number of theoretical estimates by plotting the Stanton number against the energy-thickness Reynolds number. The prediction giving the most reasonable agreement throughout the above Mach number range is that due to Fernholz (1971).The values of the skin-friction coefficient derived from velocity profiles and Preston tube data are also given.

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.

Hole Size Effect on Hemisphere Pressure Distributions

1967 ◽  
Vol 71 (676) ◽  
pp. 317-319 ◽  
Author(s):  
D. F. Morrison ◽  
L. M. Sheppard ◽  
M. J. Williams
Keyword(s):  
Atmospheric Pressure ◽  
Upper Atmosphere ◽  
Sounding Rocket ◽  
Similar Technique ◽  
Number Range ◽  
Sounding Rockets ◽  
Pressure Distributions ◽  
The Usa ◽  
Mach Numbers ◽  
Preliminary Consideration

During 1961 preliminary consideration was given by the Weapons Research Establishment to the possibility of measuring the structure of the upper atmosphere by pressure probes on sounding rockets. A similar technique had been used with success in the USA. It was proposed that velocity be measured and that an incidence meter be used to measure pitot and static pressures, from which atmospheric pressure and density could be calculated. There was available at the time comprehensive calibration information on hemispherical-headed incidence meters at Mach numbers below 3. Since the sounding rocket would be flying within the Mach number range 3 to 6 there was a need to extend the existing calibrations to higher Mach numbers. At first it seemed that the results of Baer would do this, but a closer examination showed that his results did not agree with the incidence meter calibrations.

Direct numerical simulation of hypersonic turbulent boundary layers. Part 3. Effect of Mach number

2011 ◽  
Vol 672 ◽  
pp. 245-267 ◽  
Author(s):  
L. DUAN ◽  
I. BEEKMAN ◽  
M. P. MARTÍN
Keyword(s):  
Mach Number ◽  
Boundary Layers ◽  
Large Scale ◽  
Free Stream ◽  
Turbulent Boundary Layers ◽  
Thermodynamic Quantities ◽  
Reynolds Analogy ◽  
Number Range ◽  
Free Stream Mach Number ◽  
Mach Numbers

In this paper, we perform direct numerical simulations (DNS) of turbulent boundary layers with nominal free-stream Mach number ranging from 0.3 to 12. The main objective is to assess the scalings with respect to the mean and turbulence behaviours as well as the possible breakdown of the weak compressibility hypothesis for turbulent boundary layers at high Mach numbers (M > 5). We find that many of the scaling relations, such as the van Driest transformation for mean velocity, Walz's relation, Morkovin's scaling and the strong Reynolds analogy, which are derived based on the weak compressibility hypothesis, remain valid for the range of free-stream Mach numbers considered. The explicit dilatation terms such as pressure dilatation and dilatational dissipation remain small for the present Mach number range, and the pressure–strain correlation and the anisotropy of the Reynolds stress tensor are insensitive to the free-stream Mach number. The possible effects of intrinsic compressibility are reflected by the increase in the fluctuations of thermodynamic quantities (p′rms/pw, ρ′rms/ρ, T′rms/T) and turbulence Mach numbers (Mt, M′rms), the existence of shocklets, the modification of turbulence structures (near-wall streaks and large-scale motions) and the variation in the onset of intermittency.

Techniques for Supersonic Turbojet Noise Reduction

2012 ◽  
Author(s):  
David Munday ◽  
Dan Cuppoletti ◽  
Michael Perrino ◽  
Ephraim Gutmark ◽  
Markus O. Burak ◽  
...  
Keyword(s):  
Mach Number ◽  
Noise Reduction ◽  
Secondary Flow ◽  
High Performance ◽  
Supersonic Jet ◽  
Military Aircraft ◽  
Mach Numbers

Observations and simulations are presented of a supersonic jet from a nozzle representative of high-performance military aircraft such as the Saab Gripen. The nozzle has a design Mach number of 1.56 and is examined at its design condition with a surrounding secondary flow at Mach numbers of 0.0, 0.1 and 0.3. Chevrons and internal fluidic injection by microjets each reduce the noise generated by the main jet.

Wing Section Design for Swept-Back Wings at Transonic Speeds

1957 ◽  
Vol 61 (556) ◽  
pp. 238-244 ◽  
Author(s):  
A. B. Haines
Keyword(s):  
Mach Number ◽  
Aspect Ratio ◽  
The Body ◽  
Wing Section ◽  
Pressure Distributions ◽  
Section Design ◽  
Mach Numbers ◽  
Sonic Speed ◽  
Peak Value

It is well known that the performance at high subsonic and transonic speeds of a swept-back wing-body combination in which the wing is untwisted and has the same section at all stations along the span and in which the body is not specially shaped to allow for the presence of the wing, falls far short of what would be predicted for the corresponding infinite sheared wing. For example, with a sweep of 45° and a thickness/chord ratio of 6 per cent it has been found experimentally that a rapid shock-induced increase in drag occurs above a Mach number of about 0·95 and a peak value of CD is obtained at Mach numbers slightly in excess of 1·0, whereas it can be estimated that for the corresponding infinite sheared wing, sonic speed in a direction perpendicular to the isobars (the lines joining points where the pressure is equal) would not be obtained until a Mach number of 1·18 was reached. The poorer performance of the finite swept-back wing results principally from the fact that the pressure distributions for sections near the root and tip are distorted in shape from what would be obtained on an infinite sheared wing and, as a result, the isobars tend to lose some or all of their sweep. With a moderate aspect ratio such as 3, such effects extend over most of the span at high subsonic speeds.

scholarly journals Mass Transfer Cooling on a Porous Flat Plate in Carbon-Dioxide and Air Streams

1968 ◽  
Vol 90 (4) ◽  
pp. 596-600 ◽  
Author(s):  
A. L. Laganelli ◽  
J. P. Hartnett
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Mach Number ◽  
Flat Plate ◽  
Transpiration Rate ◽  
Free Stream ◽  
Recovery Factor ◽  
Effective Length ◽  
Number Range ◽  
Porous Flat Plate

Heat transfer results are reported for a transpiration cooled porous flat plate placed in a stream of air and in a stream of CO2. The tests were performed at a Mach number of 1.96 over a range of effective length Reynolds number, from 5 million to 9.1 million, when CO2 was used as the free stream gas. A Mach number of 2.53 for an effective length Reynolds number range of 5.3 million to 8.3 million was characteristic when the free stream gas was air. The heat transfer data were normalized and presented as the ratio of the Stanton number to the no-blowing Stanton value (St/St0) as a function of the dimensionless transpiration rate F/St0. The recovery factor data were also normalized and presented as the ratio of r/r0 as a function of the transpiration rate F. The results for both the air and the CO2 free stream flows showed a reduction in heat transfer with increasing transpiration rate, using air and CO2 as the injectant gases. The measured recovery factor and the normalized recovery factor also decreased with increasing transpiration for the reported gas combinations. It was found that Rubesin’s air theory adequately predicts all of the heat transfer results including those obtained in CO2 atmospheres within the reported Mach number range. Also, the empirical theories which predict recovery factor results for air free streams can be used for air or CO2 injection into a CO2 free stream gas.

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