Monograph records
https://feeds.library.caltech.edu/people/Liepmann-H-W/monograph.rss
A Caltech Library Repository Feedhttp://www.rssboard.org/rss-specificationpython-feedgenenFri, 08 Dec 2023 12:22:33 +0000Investigations of Effects of Surface Temperature and Single Roughness Elements on Boundary-Layer Transition
https://resolver.caltech.edu/CaltechAUTHORS:LIEnacarpt890
Authors: Liepmann, Hans W.; Fila, Gertrude H.
Year: 1947
The laminar boundary layer and the position of the transition point are investigated on a heated flat plate. It was found that the Reynolds number of transition decreases as the temperature of the plate is increased. It is shown from simple qualitative analytical considerations that the effect of variable viscosity in the boundary layer due to the temperature diference produces a velocity profile with an inflection point if the wall temperature is higher than the free-stream temperature. This profile is confirmed by measurements. Furthermore, it is confirmed that, even with large deviation from the Blasius condition, the velocity and temperature profiles are very nearly identical, as predictable theoretically for a Prandtl number [sigma] of the order of 1.0 (for air, [sigma]=0.76). The instability of
injection-point profiles is discussed.
Studies of the flow in the wake of large, two-dimensional
roughness elements are presented. It is shown that a boundary laysr can separate and reattach itself to the wall without having transition take place.https://authors.library.caltech.edu/records/awb3z-zw847Investigation of Effects of Surface Temperature and Single Roughness Elements on Boundary-Layer Transition
https://resolver.caltech.edu/CaltechAUTHORS:LIEnacatn1196
Authors: Liepmann, Hans W.; Fila, Gertrude H.
Year: 1947
The laminar boundaxy layer and the position of the transition point were investigated on a heated flat plate. It was found that the Reynolds number of transition decreases as the temperature of the plate is increased. It is shown from simple qualitative analytical considerations that the effect of variable viscosity in the boundary layer due to the temperature difference produces a velocity profile with an inflection point if the wall temperature is higher than the free-stream temperature. This profile is confirmed by measurements. Furthermore, it is confirmed that even with large deviation from the Blasius condition, the velocity and temperature profiles are very nearly identical, as predictable theoretically for a Prandtl number [sigma] of the order of 1.0 (for air, [sigma] = 0.76). The instability of inflection-point profiles is discussed.
Studies of the flow in the wake of large, two-dimensional roughness elements are presented. It is shown that a boundary layer can separate and reattach itself to the wall without having transition take place.https://authors.library.caltech.edu/records/3x3eq-aj709Investigations of Free Turbulent Mixing
https://resolver.caltech.edu/CaltechAUTHORS:LIEnacatn1257
Authors: Liepmann, Hans Wolfgang; Laufer, John
Year: 1947
A discussion of the integral relations for the flow of the boundary-layer type is presented. It is shown that the characteristic laws of spread of jets, wakes, and so forth, can be obtained directly for the laminar case and, with the help of dimensional reasoning, for the turbulent case as well.
Measurements of the mean velocity, the intensity and scale of the turbulent fluctuations, and of the turbulent shear in a two-dimensional mixing zone are presented. The results of these measurements are compared with the mixing-length theories. It is shown that both mixing length and exchange coefficient vary across the mixing zone. The theories based on the assumption of constant mixing length or exchange coefficient are thus in error.
A discussion of the energy balance of the fluctuating motion is given and the triple point correlation is estimated.https://authors.library.caltech.edu/records/ebd4r-t6m96On the Spectrum of Isotropic Turbulence
https://resolver.caltech.edu/CaltechAUTHORS:LIEnacatn2473
Authors: Liepmann, H. W.; Laufer, J.; Liepmann, Kate
Year: 1951
Measurements of the spectrum and correlation functions at large Reynolds number (RN ~ 10^5 based on the grid mesh) have been made, as well as a series of accurate spectrum measurements at lower Reynolds number (RN ~ 10^4).
The results are compared with the theoretical laws proposed in recent years. It is found that the measurements at large Reynolds numbers exhibit a range of frequencies where the spectrum is nearly of the form n^- 5/3.
The largest part of the spectrum in the initial stage of decay at the lower Reynolds number was found to follow closely the simple spectrum A/[B + n^2] , where A and B are constants and n is the frequency of fluctuation. At x/M = 1000 (where x is the distance behind the grid and M is the mesh size) the spectrum approaches a Gaussian distribution.
The second, fourth, and sixth moments of the spectrum have been computed from the measurements and are discussed In relation to theoretical results.
The significance of the number of zeros of the fluctuating velocity u(t) is discussed and examples of measurements for the determination of the microscale of turbulence [lambda] from zero counts are given.https://authors.library.caltech.edu/records/58rp5-bsf15On Reflection of Shock Waves from Boundary Layers
https://resolver.caltech.edu/CaltechAUTHORS:LIEnacarpt1100
Authors: Liepmann, H. W.; Roshko, A.; Dhawan, S.
Year: 1952
Measurements of the reflection characteristics of shock waves from a flat surface with a laminar and turbulent boundary layer are presented. The investigations were carried out at Mach numbers from about 1.3 to 1.5 and a Reynolds number of 0.9 x 10^4.
THe difference in the shock-wave interaction with laminar and turbulent boundary layers, first found in transonic flow is confirmed and ,investigated in detail for supersonic flow. The relative upstream influence of a shock wave impinging on a given boundary layer has been measured for both laminar and turbulent layers. The upstream influence of a shock wave in the laminar layer is found to be of the order of 50 bounday-layer thicknesses as compared with about 5 in the turbulent case. Separation almost always occurs in the laminar boundary layer. The separation is restricted to a region of finite extent upstream of the the shock wave. In the turbulent case no separation was found. A model of the flow near the point of impingement of the shock wave on the boundary layer is given for both cases. The difference between impulse-type and step-type shock waves is discussed and their interactions with the boundary layer are compared.
Some general considerations on the experimental production of shock waves from wedges and cones are presented, as well as a discussion of boundary layer in supersonic flow. A few exampies of reflection of shock waves from supersonic shear layers are also presented.https://authors.library.caltech.edu/records/2haej-rzc24Counting Methods and Equipment for Mean-Value Measurements in Turbulence Research
https://resolver.caltech.edu/CaltechAUTHORS:LIEnacatn3037
Authors: Liepmann, H. W.; Robinson, M. S.
Year: 1953
This report deals with methods of measuring the probability distributions and mean values of random functions as encountered in turbulence research. Applications to the measurement of probability distributions of the axial velocity fluctuation u(t) and its derivative du/dt in isotropic turbulence are shown. The assumption of independent probabilities of u(t) and du/dt, which has been used as an approximation in the application of zero counts to the measurement of the microscale of turbulence [lambda], is investigated. The results indicate that the assmuption is satisfied within a few percent and that there is, so far, no evidence that the systematic difference between [lambda] measured from zero counts and [lambda] measured independently can be traced entirely to the statistical dependence of u and du/dt.
The chronological development of apparatus is described, concluding with the present 10-channel statistical analyzer based upon a system of pulse amplitude modulation followed by an amplitude discriminator and a counter. A discussion of the relative merits of various systems is included to indicate the reasons for this choice.https://authors.library.caltech.edu/records/dgqbv-ydk93Shearing-Stress Measurements by use of a Heated Element
https://resolver.caltech.edu/CaltechAUTHORS:LIEnacatn3268
Authors: Liepmann, H. W.; Skinner, G. T.
Year: 1954
The rate of local heat transfer from a solid surface to a moving fluid is related to the local skin frinction. Measurements of the heat transmission from small elements embedded in the surface of a solid can thus be used to botain local skin-friction coefficients. This method was applied by Fage and Falkner for laminar boundary layers and by Ludwieg for turbulent boundary layers. The present report discussed the possible range of application of such an instrument in low- and high-speed flow and presents experimental data to show that a very simple instrument can be used to obtain laminar and turbulent skin-friction coefficients with a single calibration. The instrument consists of an ordinary hot-wire cemented into a groove in the surface. The heat loss from the wire is proportional to the cube root of the wall shearing stress, and the constant of proportionality may be found by one calibration, for example, in laminar flow.https://authors.library.caltech.edu/records/z1sfr-x9a98On the Contribution of Turbulent Boundary Layers to the Noise inside a Fuselage
https://resolver.caltech.edu/CaltechAUTHORS:CORnacatm1420
Authors: Corcos, G. M.; Liepmann, H. W.
Year: 1956
The following report deals i preliminary fashion with the transmission through a fuselage of random noise generated on the fuselage skin by a turbulent boundary layer. The concept of attenuation is abandoned and instead the problem is formulated as a sequence of two linear couplings: the turbulent boundary layer fluctuations excite the fuselage skin in lateral vibrations and the skin vibrations induce sound inside the fuselage. The techniques used are those required to determine the response of linear systems to random forcing functions of several variables. A certain degree of idealization has been resorted to. Thus the boundary layer is assumed locally homogeneous, the fuselage skin is assumed flat, unlined and free from axial loads and the "cabin" air is bounded only by the vibrating plate so that only outgoing waves are considered. Some of the details of the statistical description have been simplified in order to reveal the basic features of the problem.
The results, strictly applicable only to the limiting case of thin boundary layers, show that the sound pressure intensity is proportional to the square of the free stream density, the square of cabin air density and inversely proportional to the first power of the damping constant and to the second power of the plate density. The dependence on free stream velocity and boundary layer thickness cannot be given in general without a detailed knowledge of the characteristics of the pressure fluctuations in the boundary layer (in particular the frequency spectrum). For a flat spectrum the noise intensity depends on the fifth power of the velocity and the first power of the boundary layer thickness. This suggests that boundary layer removal is probably not an economical means of decreasing cabin noise.
In general, the analysis presented here only reduces the determination of cabin noise intensity to the measurement of the effect of any one of four variables (free stream velocity, boundary layer thiclkness, plate thickness or the characteristic velocity of propagation in the plate).
The plate generates noise by vibrating in resonance over a wide range of frequencies and increasing the damping constant is consequently an effective method of decreasing noise generation.
One of the main features of the results is that the relevent quantities upon which noise intensity depends are non-dimensional numbers in which boundary layer and plate properties enter as ratios. This is taken as an indication that in testing models of structures for boundary layer noise it is not sufficient to duplicate in the model the structural characteristics of the fuselage. One must match properly the characteristics of the exciting pressure fluctuations to that of the structure.https://authors.library.caltech.edu/records/3q4c9-e8r53On the acoustic radiation from boundary layers and jets
https://resolver.caltech.edu/CaltechAUTHORS:LIEgalcit54
Authors: Liepmann, H. W.
Year: 2008
In the following a general discussion of aerodynamically created sound is given. The study is essentially theoretical in nature, but arrives at a description of the physical phenomena in such a fashion as to yield an immediate access to experiments.
First, the problem of aerodynamic noise is defined and two simple mechanical analogues discussed. Then, the general equations of motion of a viscous, compressible fluid are rearranged in a form suitable for a comparison with Lighthill's approach. However, this approach is not being followed through. Instead, the concept of induced velocities due to displacement effects is put forward and carried through for noise produced by boundary layer flow. The same concept is then extended to describe the sound field created by a jet.https://authors.library.caltech.edu/records/7hmt5-ctr70Investigations of the Interaction of Boundary Layer and Shock Waves in Transonic Flow
https://resolver.caltech.edu/CaltechAUTHORS:20151111-142642000
Authors: Liepmann, H. W.
Year: 2015
A discussion is given of the interaction between shock waves and boundary layer and of the formation of shocks in transonic flow, based on measurements of transonic flow past a 12% thick circular arc profile. It is found that:
a. The shock wave pattern at a given Mach number can be completely altered by changing the boundary layer.
b. Shock waves can interact with a boundary layer in a manner similar to the reflection of a wave from a free jet boundary. Shock waves do not necessarily cause boundary layer separation.
c. There exist two types of possible transonic flow past a given symmetrical boundary. One symmetrical about the maximum thickness point and one asymmetrical about this point. The first can be identified with the known symmetrical potential solutions. In this case, recompression begins without a shock wave. The asymmetrical case is characterized by an expansion of the flow up to the shock wave. The shock wave in the symmetrical case is related to the "limiting" line of potential theory, the shock wave in the asymmetrical case to the shock waves occurring in de Laval nozzles.https://authors.library.caltech.edu/records/90t28-1zn74