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Flight experiments were conducted on a single-engine turboprop aircraft fitted with a 92-in-chord, 3-ft-span natural laminar flow glove at glove section lift coefficients from 0.15 to 1.10. The boundary-layer transition measurement methods used included sublimating chemicals and surface hot-film sensors. Transition occurred downstream of the minimum pressure point. Hot-film sensors provided a well-defined indication of laminar, laminar-separation, transitional, and turbulent boundary layers. Theoretical calculations of the boundary-layer parameters provided close agreement between the predicted laminar-separation point and the measured transition location. Tollmien-Schlichting (T-S) wave growth n-factors between 15 and 17 were calculated at the predicted point of laminar separation. These results suggest that for many practical airplane cruise conditions, laminar separation (as opposed to T-S instability) is the major cause of transition in predominantly two-dimensional flows.

Using a perturbation approach, we make rigorous the formal boundary layer asymptotic analysis of Turcotte, Spence and Bau from the early eighties for the vertical flow of an internally heated Boussinesq fluid in a vertical channel with viscous dissipation and pressure work. A key point in our proof is to establish the non-degeneracy of a special solution of the Painlev\'{e}-I transcendent. To this end, we relate this problem to recent studies for the ground states of the focusing nonlinear Schr\"{o}dinger equation in an annulus. We also relate our result to a particular case of the well known Lazer-McKenna conjecture from nonlinear analysis.

In this paper we extend the ideas presented in Onofrei and Vernescu [\textit{Asymptotic Analysis, 54, 2007, 103-123}] and introduce suitable second order boundary layer correctors, to study the $H^1$-norm error estimate for the classical problem in homogenization. Previous second order boundary layer results assume either smooth enough coefficients (which is equivalent to assuming smooth enough correctors $\chi_j,\chi_{ij}\in W^{1,\infty}$), or smooth homogenized solution $u_0$, to obtain an estimate of order $\displaystyle O(\epsilon^{\frac{3}{2}})$. For this we use the periodic unfolding method developed by Cioranescu, Damlamian and Griso [\textit{C. R. Acad. Sci. Paris, Ser. I 335, 2002, 99-104}]. We prove that in two dimensions, for nonsmooth coefficients and general data, one obtains an estimate of order $\displaystyle O(\epsilon^\frac{3}{2})$. In three dimenssions the same estimate is obtained assuming $\chi_j,\chi_{ij}\in W^{1,p}$, with $p>3$. We also discuss how our results extend, in the case of nonsmooth coefficients, the convergence proof for the finite element multiscale method proposed by T.Hou et al. [\textit{ J. of Comp. Phys., 134, 1997, 169-189}] and the first order correctoranalysis for the first eigenvalue of a composite media obtained by Vogelius et al.[\textit{Proc. Royal Soc. Edinburgh, 127A, 1997, 1263-1299}].

A supplement is presented for a turbulent boundary layer nozzle analysis computer program. It describes the program calculation sequence and presents a detailed documentation of each subroutine. Important equations are derived explicitly, and improvements to the program are discussed.

A computation procedure is described for predicting flow fields which develop when successive interactions between oblique shock waves and a turbulent boundary layer occur. Computations were carried out for axisymmetric internal flows at free stream Mach numbers = 3.82 and 2.82. The effect of boundary layer bleed was considered for the free stream Mach number = 2.82 flow. A control volume analysis is used to predict changes in the flow field across the interactions. Two bleed flow models were considered. A turbulent boundary layer program was used to compute changes in the boundary layer between the interactions. The results given are for flows with two shock wave interactions and for bleed at the second interaction site. In principle the method described may be extended to account for additional interactions. The predicted results are compared with measured results and are shown to be in good agreement when the bleed flow rate is low (on the order of 3 percent of the boundary layer mass flow), or when there is no bleed. As the bleed flow rate is increased, differences between the predicted and measured results become larger. Shortcomings of the bleed flow models at higher bleed flow rates are discussed.

As both computers and flow-analyzing equations have increased in sophistication, Computational Fluid Dynamics (CFD) has evolved into a fixture for advanced aircraft design. While CFD codes have improved in accuracy and efficiency, their ability to encompass viscous effects is lacking in certain areas. For example, current CFD codes cannot accurately predict or correct for the increased drag due to these viscous effects at some flow conditions. However, by analyzing an airfoil's turbulent boundary layer, one can predict not only flow separation via the shape factor parameter, but also viscous drag via the momentum thickness. Various codes have been written which can calculate turbulent boundary layer parameters. The goal of my research is to develop procedures for modifying an airfoil (via its local pressure distribution) to eliminate boundary layer separation and/or to reduce viscous drag. The modifications to the local pressure distribution necessary to achieve these objectives will be determined using a direct-iterative method installed into a turbulent boundary layer analyzer. Furthermore, the modifications should preserve the basic characteristics of the original airfoil.

The general objective of the present work is to explore the use of Rapid Distortion Theory (RDT) in analysis of the two-point statistics of the log-layer. RDT is applicable only to unsteady flows where the non-linear turbulence-turbulence interaction can be neglected in comparison to linear turbulence-mean interactions. Here we propose to use RDT to examine the structure of the large energy-containing scales and their interaction with the mean flow in the log-region. The contents of the work are twofold: First, two-point analysis methods will be used to derive the law-of-the-wall for the special case of zero mean pressure gradient. The basic assumptions needed are one-dimensionality in the mean flow and homogeneity of the fluctuations. It will be shown that a formal solution of the two-point correlation equation can be obtained as a power series in the von Karman constant, known to be on the order of 0.4. In the second part, a detailed analysis of the two-point correlation function in the log-layer will be given. The fundamental set of equations and a functional relation for the two-point correlation function will be derived. An asymptotic expansion procedure will be used in the log-layer to match Kolmogorov's universal range and the one-point correlations to the inviscid outer region valid for large correlation distances.

The general objective of the present work is to explore the use of Rapid Distortion Theory (RDT) in analysis of the two-point statistics of the log-layer. RDT is applicable only to unsteady flows where the non-linear turbulence-turbulence interaction can be neglected in comparison to linear turbulence-mean interactions. Here we propose to use RDT to examine the structure of the large energy-containing scales and their interaction with the mean flow in the log-region. The contents of the work are twofold: First, two-point analysis methods will be used to derive the law-of-the-wall for the special case of zero mean pressure gradient. The basic assumptions needed are one-dimensionality in the mean flow and homogeneity of the fluctuations. It will be shown that a formal solution of the two-point correlation equation can be obtained as a power series in the von Karman constant, known to be on the order of 0.4. In the second part, a detailed analysis of the two-point correlation function in the log-layer will be given. The fundamental set of equations and a functional relation for the two-point correlation function will be derived. An asymptotic expansion procedure will be used in the log-layer to match Kolmogorov's universal range and the one-point correlations to the inviscid outer region valid for large correlation distances.

The general objective of the present work is to explore the use of Rapid Distortion Theory (RDT) in analysis of the two-point statistics of the log-layer. RDT is applicable only to unsteady flows where the non-linear turbulence-turbulence interaction can be neglected in comparison to linear turbulence-mean interactions. Here we propose to use RDT to examine the structure of the large energy-containing scales and their interaction with the mean flow in the log-region. The contents of the work are twofold: First, two-point analysis methods will be used to derive the law-of-the-wall for the special case of zero mean pressure gradient. The basic assumptions needed are one-dimensionality in the mean flow and homogeneity of the fluctuations. It will be shown that a formal solution of the two-point correlation equation can be obtained as a power series in the von Karman constant, known to be on the order of 0.4. In the second part, a detailed analysis of the two-point correlation function in the log-layer will be given. The fundamental set of equations and a functional relation for the two-point correlation function will be derived. An asymptotic expansion procedure will be used in the log-layer to match Kolmogorov's universal range and the one-point correlations to the inviscid outer region valid for large correlation distances.

Boundary-layer stability analyses of mean flows extracted from unstructured-grid Navier- Stokes solutions have been performed. A procedure has been developed to extract mean flow profiles from the FUN3D unstructured-grid solutions. Extensive code-to-code validations have been performed by comparing the extracted mean ows as well as the corresponding stability characteristics to the predictions based on structured-grid solutions. Comparisons are made on a range of problems from a simple at plate to a full aircraft configuration-a modified Gulfstream-III with a natural laminar flow glove. The future aim of the project is to extend the adjoint-based design capability in FUN3D to include natural laminar flow and laminar flow control by integrating it with boundary-layer stability analysis codes, such as LASTRAC.

The stability characteristics of an uncooled sharp cone at zero angle of attack have been calculated at an edge Mach number of 6.8 and a stagnation temperature of 728 K (1310 R) using spatial normal modes. Transformed flat-plate boundary-layer profiles were used for the mean flow. By a consideration of both two-dimensional and oblique waves, it has been determined that instability at this Mach number is primarily a two-dimensional second-mode instability. Comparisons with available stability experiments show that the calculated unstable frequencies were present in the experiments, but the measured amplification rates differ from the calculated. A band of unstable high frequencies above the second-mode region that starts at about R=1400 is not accounted for by the theory. Comparisons of amplitude-growth curves for both cones and flat plates at March numbers 4.5, 5.8 and 6.8 show that the most unstable normal modes on a cone have at least as much total growth as on a flat plate at Mach 4.5 and more total growth at the other two Mach numbers. Keywords: Boundary layer stability; Hypersonic linear stability theory; Hypersonic cone stability.

For the generation of impulsive sound waves caused by parallel interaction of a vortex and an airfoil in a plane flow field, two different mechanisms are responsible by experimental evidence. The first one originates from the area of the stagnation point of the airfoil: a temporal increase of pressure and density - in consequence of the incoming vortex - relaxes by sound wave emission, when the vortex vanishes behind the airfoils nose. This is called a compressibility wave. The second one is reasoned by a supersonic flow regime, which appears, when the stationary airfoil flow is augmented by the flow field of the vortex: at the shoulder of the airfoil we get an unsteady return to subsonics by a shock wave. This moves upstream after the vortex has passed and is named transonic wave. Evidently both mechanisms only occur, if the flow field at the airfoil is augmented by the vortex, i.e. the vortex has a special spin orientation with respect to the airfoil. Keywords: Sound generation mechanism; Vortex airfoil interaction; Image processing; Image analysis; Vortex generation system.

This paper presents the results of the analysis of Mach 8.0 flow over a seven degree half-angle cone. The purpose of this analysis was to develop techniques to examine boundary layer transition at hypersonic velocities. The specific objectives were to look for second mode instability waves characteristic of the transition process and to quantify the percentage of turbulent flow. Two sets of data were used in this analysis. The first set of data was taken at several axial positions at a freestream Reynolds number 4.265 million per meter. This data was used to develop the analysis techniques. The second set of data was taken at station 35 for Reynolds numbers of 3.28, 3.94,4.92, and 6.56 million per meter. Spectral analysis was used to identify 2nd mode disturbances, if they existed. The energy associated with the disturbances was then removed from the data signal to produce a new signal. The new signal was then evaluated using conditional sampling techniques. Additional methods used to assess turbulent intermittency were histogram analysis and examination of the power spectrum of the data signal. It was determined that removal of the disturbances from the raw data signal produced a cleaner signal. However, the new signals were not amenable to conditional sampling techniques. The histogram analysis proved to be inconclusive. Examination of the power spectrum showed that a laminar flow could be identified by the presence of a strong peak corresponding to the 2nd mode disturbances, but could not be used to identify a flow as being turbulent by the absence of this peak.

A critical analysis is presented of the available wind tunnel data simulating nosetip boundary layer transition on reentry vehicles. It is agreed that transition should depend on surface roughness, surface temperature, and surface curvature. The Reynolds number based on momentum thickness at the transition location (Re sub theta) is used to measure transition, and the surface roughness effect is described by the Reynolds number Rek,k based on roughness height and conditions at the tops of the roughness elements. The wall temperature dependence may be removed by use of the kinematic viscosity in the 'middle' of the boundary layer to compute Re sub theta. The available data suggest a dependence on theta/Rc (Rc = surface radius of curvature), which could be due to centrifugal acceleration or pressure gradient. Accounting for this curvature effect appears to eliminate the need for a separate transition 'onset' criterion. A tentative correlation is presented, and the limitations of the existing data are discussed.

This paper analyzes the impact dispersion of a ballistic reentry vehicle caused by the transitory lift and drag components encountered during the brief altitude regime where the vehicle's boundary layer is changing from laminar to turbulent behavior. A simple physical model of the phenomenon is developed using adjustable parameters since the basic aerothermal materials physics is poorly understood. The adjustable parameters are identifiable with physical assumptions and one result is to focus attention on the nature of these assumptions in order to guide further theoretical and experimental work which can develop a better understanding of the phenomenon, particularly its impact on materials and structures technology. Under the assumptions made, numerical estimates are made of the boundary layer transition dispersion for a practical range of vehicle and trajectory parameters, and then compared with contributions from density, wind velocity, and drag variations and from the roll-through-zero effect. The boundary layer effects are found to be too large to be ignored and further studies are needed to bound the behavior. Additional studies of the Materials and Structures Technology Base Programs related reentry vehicle mission requirements will be made in a subsequent paper. (Author)

An analysis is developed for the boundary layer flows, with dissociated and ionized air species in chemical equilibrium or with finite rate chemistry, about general body geometries at zero angle of attack. The partial differential equations for the boundary layer, about axisymmetric and two- dimensional bodies, are transformed to the von Mises stream function coordinate. The resulting equations are reduced to an algebraic system obtained by adopting an explicit finite-difference method of solution. These equations provide the basis for a computer program to obtain numerical solutions. Laminar and several turbulent viscosity models are included in the analysis. Turbulent diffusivity models based upon the law of the wall, law of the wake, a curve-fit to pipe flow eddy viscosity and laminar viscosity contributions are introduced.

Atmospheric precipitation forecasting has long been a difficult problem. One of the main reasons is possibly that the median scale triggering regime of precipitation has not been sufficiently explained. This paper uses observation data at the 325-m meteorological tower at the Inst. of Atmospheric Physics of the Chinese Academy of Sciences, sounding data of acoustic radar, and conventional data to analyze the triggering regime of two heavy rainfalls in the Beijing area from 0230 to 0430 hours, 27 Aug. 1980. By analyzing the continuous observation data through sounding of the boundary layer, it was discovered that the heavy precipitation process of the large scale front precipitation is triggered by a medium scale system. In this process, the medium scale system with period about 8 to 9 minutes was the one. This wavy motion occurred at night, caused by shallow density flow at low layers; this is a gravity internal wave. In the Beijing area when a front passes at night, consideration should be given to the superposition of mountain wind and slightly deviated northerly air flow; this superposition intensifies the density flow at low layers following the front and will cause the establishment of the 2nd temperature inversion layer and stimulate the inversion layer to produce gravity waves. Thus, for night passage of a weather front, consideration should be given to this amplitude increase function. Chinese translations.

This paper presents the linear Global stability analysis of the incompressible axisymmetric boundary layer on a circular cone. The base flow is considered parallel to the axis of cone at the inlet. The angle of attack is zero and hence the base flow is axisymmetric. The favorable pressure gradient develops in the stream-wise direction due to cone angle. The Reynolds number is calculated based on the cone radius (a) at the inlet and free-stream velocity ($U_{\infty}$). The base flow velocity profile is fully non-parallel and non-similar. Linearized Navier-Stokes equations (LNS) are derived for the disturbance flow quantities in the spherical coordinates. The LNS are discretized using Chebyshev spectral collocation method. The discretized LNS along with the homogeneous boundary conditions forms a general eigenvalues problem. Arnoldi's iterative algorithm is used for the numerical solution of the general eigenvalues problem. The Global temporal modes are computed for the range of Reynolds number from 174 to 1046, semi-cone angles $2^o$, $4^o$, $6^o$ and azimuthal wave numbers from 0 to 5. It is found that the Global modes are more stable at higher semi-cone angle $\alpha$, due to the development of favorable pressure gradient. The effect of transverse curvature is reduced at higher semi-cone angles ($\alpha$). The spatial structure of the eigenmodes show that the flow is convectively unstable. The spatial growth rate ($A_x$) increases with the increase in semi-cone angle ($\alpha$) from $2^o$ to $6^o$. Thus, the effect of increase in semi-cone angle ($\alpha$) is to reduces the temporal growth rate ($\omega_i$) and to increases the spatial growth rate ($A_x$) of the Global modes at a given Reynolds number.

This report summarizes results of stability analyses performed for a slender sharp cone with and without injection of air through a porous strip. The free-stream parameters are related to the experiments in the Caltech s T5 shock tunnel. The analysis is focused on pure aerodynamic effects in the framework of perfect gas model. It was shown that the injection leads to destabilization of acoustic modes in the near-field relaxation region. To reduce this detrimental effect it was suggested to alter the injector surface shape or use suction-blowing of zero net injection. However, stability computations showed that these modifications did not improve the injector performance. The porous wall effect on the acoustic instability was also examined. It was found that porosity stabilizes the boundary layer flow with and without injection. This indicates that the injector performance can be improved, if the porous layer, which is used for the injection, is protruded downstream to cover the near-filed relaxation region.

The viscous flow phenomena associated with sharp and blunt trailing edge airfoils were investigated. Experimental measurements were obtained for a 17 percent thick, high performance GAW-1 airfoil. Experimental measurements consist of velocity and static pressure profiles which were obtained by the use of forward and reverse total pressure probes and disc type static pressure probes over the surface and in the wake of sharp and blunt trailing edge airfoils. Measurements of the upper surface boundary layer were obtained in both the attached and separated flow regions. In addition, static pressure data were acquired, and skin friction on the airfoil upper surface was measured with a specially constructed device. Comparison of the viscous flow data with data previously obtained elsewhere indicates reasonable agreement in the attached flow region. In the separated flow region, considerable differences exist between these two sets of measurements.

Coking reactions of pyrolysis gas in charring ablation material

Application of a quasi-3D approach to the aerodynamic analysis of several radial turbine configurations is described. The objective was to improve the rotor aerodynamic characteristics by hub-shroud contouring. The approach relies on available 2D inviscid methods coupled with boundary layer analysis to calculate profile, mixing, and endwall losses. Windage, tip clearance, incidence, and secondary flow losses are estimated from correlations. To eliminate separation along the hub and blade suction surfaces of a baseline rotor, the analysis was also applied to three alternate hub-shroud geometries. Emphasis was on elimination an inducer velocity overshoot as well as increasing hub velocities. While separation was never eliminated, the extent of the separated area was progressively reduced. Results are presented in terms of mid-channel and blade surface velocities; kinetic energy loss coefficients; and efficiency. The calculation demonstrates a first step for a systematic approach to radial turbine design that can be used to identify and control aerodynamic characteristics that ultimately determine heat transfer and component life. Experimentation will be required to assess the extent to which flow and boundary layer behavior were predicted correctly.

Nitrogen injection effects on boundary layer of graphite surface in high-temperature dissociated airflow

Computer program for finite difference equation analysis of in-depth response of materials exposed to high temperature environment

An axisymmetric panel code and a three dimensional Navier-Stokes code (used as an inviscid Euler code) were verified for low speed, high angle of attack flow conditions. A three dimensional Navier-Stokes code (used as an inviscid code), and an axisymmetric Navier-Stokes code (used as both viscous and inviscid code) were also assessed for high Mach number cruise conditions. The boundary layer calculations were made by using the results from the panel code or Euler calculation. The panel method can predict the internal surface pressure distributions very well if no shock exists. However, only Euler and Navier-Stokes calculations can provide a good prediction of the surface static pressure distribution including the pressure rise across the shock. Because of the high CPU time required for a three dimensional Navier-Stokes calculation, only the axisymmetric Navier-Stokes calculation was considered at cruise conditions. The use of suction and tangential blowing boundary layer control to eliminate the flow separation on the internal surface was demonstrated for low free stream Mach number and high angle of attack cases. The calculation also shows that transition from laminar flow to turbulent flow on the external cowl surface can be delayed by using suction boundary layer control at cruise flow conditions. The results were compared with experimental data where possible.

A computer aided boundary layer analysis, which draws on available theoretical and empirical data, is presented. The analysis allows boundary layer data to be extracted from measurements influenced by a normal pressure gradient, comparison of boundary layer measurements against standard laminar and turbulent profiles, and evaluation of important boundary layer parameters. This technique is illustrated by example.

This paper describes the application of non-asymptotic triple-deck theory of transonic shock-turbulent boundary layer interaction as a local modular element in the global calculation of supercritical airfoil flow fields. (Author)

The structure of an incompressible turbulent boundary layer subjected to a strong adverse pressure gradient is studied by means of an asymptotic analysis of the Reynolds time-averaged equations. Limit-process expansions developed in the limit of large Reynolds number reveal a relatively thick nondefect layer in the outer region of the boundary layer near the exterior inviscid flow, and a relatively thin layer near the wall. To leading orders of approximation, the momentum balance involves convection, pressure gradient, and turbulent stress in the outer layer, and pressure gradient, and turbulent and viscous stresses in the inner layer. The asymptotic expansions for these two layers are matched in an arbitrary intermediate region, wherein the streamwise velocity has a square-root dependence and the Reynolds stress has a corresponding linear dependence on the normal coordinate. The leading-order approximations for the outer and inner layers give rise to similarity formulations, from which appropriate similarity formulations for the distinguished intermediate layer have been identified and developed. These latter formulations are employed to analyze available experimental data.

The structure of an incompressible turbulent boundary layer subjected to a strong adverse pressure gradient is studied by means of an asymptotic analysis of the Reynolds time-averaged equations. Limit-process expansions developed in the limit of large Reynolds number reveal a relatively thick nondefect layer in the outer region of the boundary layer near the exterior inviscid flow, and a relatively thin layer near the wall. To leading orders of approximation, the momentum balance involves convection, pressure gradient, and turbulent stress in the outer layer, and pressure gradient, and turbulent and viscous stresses in the inner layer. The asymptotic expansions for these two layers are matched in an arbitrary intermediate region, wherein the streamwise velocity has a square-root dependence and the Reynolds stress has a corresponding linear dependence on the normal coordinate. The leading-order approximations for the outer and inner layers give rise to similarity formulations, from which appropriate similarity formulations for the distinguished intermediate layer have been identified and developed. These latter formulations are employed to analyze available experimental data.

The structure of an incompressible turbulent boundary layer subjected to a strong adverse pressure gradient is studied by means of an asymptotic analysis of the Reynolds time-averaged equations. Limit-process expansions developed in the limit of large Reynolds number reveal a relatively thick nondefect layer in the outer region of the boundary layer near the exterior inviscid flow, and a relatively thin layer near the wall. To leading orders of approximation, the momentum balance involves convection, pressure gradient, and turbulent stress in the outer layer, and pressure gradient, and turbulent and viscous stresses in the inner layer. The asymptotic expansions for these two layers are matched in an arbitrary intermediate region, wherein the streamwise velocity has a square-root dependence and the Reynolds stress has a corresponding linear dependence on the normal coordinate. The leading-order approximations for the outer and inner layers give rise to similarity formulations, from which appropriate similarity formulations for the distinguished intermediate layer have been identified and developed. These latter formulations are employed to analyze available experimental data.

In biology, the importance of fluid drag, diffusion, and heat transfer both internally and eternally, suggest the boundary layer as an important subject of the investigation the complexities of biological systems present significant and unique challenges to analysis by experimental fluid dynamics.

This thesis evaluates the ability of the IRTOOL computer simulation program to predict mirages. Using identical input conditions taken from the MAPTIP experiment database, predicted Minimum Mirage Range (MMR) and Maximum Intervision Range (MIVR) from both the IRTOOL and IRBLEM models were extracted and compared with the measurements recorded in the database. By comparison of the algorithms it was found that discrepancies in IRTOOL mirage prediction could be ascribed to the input function for significant ocean wave height, which gave values much greater than measured or used in IRBLEM. For a significant wave height close to the measured value the IRTOOL predictions were in very close agreement with observation and with IRBLEM. IRTOOL predictions were in all cases within 2.7 km and in most cases within 1.3 km of the measurements for all ranges varying from about 7-26 km. The strong temperature gradient predicted by the model within a few meters of the water surface, uncertainties in the measured range, and the variation of 0.8 to 2 deg C in Air Sea Temperature Difference are sufficient to account for the observed deviations. Differences between the model predictions and some of the problems encountered are also discussed.

A performance analysis has been made to determine whether boundary-layer control by suction might reduce the minimum take-off distance of a four-place or five-place liaison-type airplane below that obtainable with conventional high-lift devices. The airplane was assumed to be capable of operating from airstrips having a ground friction coefficient of 0.2. The pay load was fixed at 1500 pounds and the wing span was varied from 30 to 100 feet, the aspect ratio from 5 to 15, and the power from 200 to 1300 horsepower. Maximum lift coefficients of 5.0 and 2.8 were assumed for the airplanes with and without boundary-layer control, respectively. A conservative estimate of the boundary-layer-control-equipment weight was included. The effects of the boundary-layer control on total take-off distance, ground run, and stalling speed were determined The analysis indicates that the addition of boundary-layer control does not reduce the absolute minimum total take-off distance that is obtained with an airplane having a low wing loading and a moderately low aspect ratio. The effectiveness of boundary-layer control in reducing the total take-off distance for a given maximum speed improves with increasing aspect ratio and, for wing loadings of 10 pounds per square foot or more and an aspect ratio of 10 or more, the addition of boundary-layer control results in a decrease in the total take-off distance. For a given maximum speed the ground run was reduced for all configurations by the addition of boundary-layer control. The reduction was negligible for aspect ratio of 5 but was 10 to 30 percent for aspect ratios of 10 and 15. The stalling speed for a given maximum speed was reduced 20 to 25 percent for all configurations by application of boundary-layer control. A reduction in the weight of the boundary-layer control equipment would result in an appreciable decrease in the total take-off and ground run distances, but would have a negligible effect on the stalling speed.

Transmission and reception of electromagnetic energy \"by communications, weapons, and active/passive sensor systems is known to be strongly influenced by an atmospheric phenomena known as ducting, caused by refractive layers in the atmosphere of marine environments. The Naval Postgraduate School (NPS) has developed a Marine Atmospheric Boundary Layer (MABL) model which can be used to predict, over a 24 hour period, the refractive profile of the lower atmosphere. This thesis examines the model from the statistical/trend analysis approach to examine whether the model can be used as a valid predictor of refractive/ducting conditions.

This report presents results of wake survey and boundary layer profile measurements for a model representing a twin-screw displacement ship. These measurements were obtained in a wind tunnel using hot wire anemometers. Wake surveys were conducted in the propeller disk and ahead of the propeller plane with and without the propeller operating. Circumferential distributions of the longitudinal, tangential, and radial velocity components at four radial locations and a harmonic analysis of each component are included. The boundary layer profile measurements were obtained at three longitudinal locations in the area of the shafting and struts. A comparison of full-scale, model-scale towing tank, and wind tunnel measurements is presented. Model-scale boundary layer velocity profiles ahead and astern of the propeller plane were smaller than the full-scale profiles. The model wake surveys ahead of the propeller plane showed no significant differences with and without the propeller operating. The data from the wind tunnel wake survey in the propeller plane agreed better with the full-scale data than the towing tank data.

In a previous report, the authors analyzed the stability of the lower branch solutions of the incompressible (M sub infinity = 0) Falkner-Skan boundary layers. This report extends this work by including the effect of Mach number M sub infinity on the stability of adiabatic case. (Author)

Correlation analysis of windspeed in the Redstone Arsenal, Alabama boundary-layer along with the variation of the exponent in the power-law wind profile for concurrent tower, LDV, and pibal data are presented for 36 cases between 13 January and 1 April 1986. Wind profiles were available up to 90 m, 450 m, and 915 m for the tower, LDV, and pibal data, respectively. The correlation between the scalar wind at 4 m with the scalar wind at 90 m is .84 for the tower data improving to .92 when the scalar wind at 19 m is correlated with the scalar wind at 90 m. The correlations derived from the LDV data set also showed the correlations improving as the reference point is raised, but the correlations were lower and dispersed over a wider range.

The present report discusses the development of a time-dependent Navier-Stokes code for use in predicting transonic shock-wave boundary layer interaction. In addition, various test cases which have been performed are discussed. The algorithm used to solve the equations is based upon the consistently split linearized block implicit method of Briley and McDonald (15, 16). The philosophy and use of a solution adaptive mesh is also described. The test cases studied in this report are: normal shock wave boundary layer interaction in a constant area circular pipe, steady transonic flow over an axisymmetric bump, unsteady flow over an axisymmetric bump, supersonic flow over an axisymmetric compression corner and oblique shock wave impingement on a flat plate boundary layer. (Author)

The present report discusses the development of a time-dependent Navier-Stokes code for use in predicting transonic shock-wave boundary layer interaction. In addition, various test cases which have been performed are discussed. The algorithm used to solve the equations is based upon the consistently split linearized block implicit method of Briley and McDonald (15, 16). The philosophy and use of a solution adaptive mesh is also described. The test cases studied in this report are: normal shock wave boundary layer interaction in a constant area circular pipe, steady transonic flow over an axisymmetric bump, unsteady flow over an axisymmetric bump, supersonic flow over an axisymmetric compression corner and oblique shock wave impingement on a flat plate boundary layer. (Author)

An analytical study of linear-amplifying instabilities of a laminar boundary layer as found in the experimental data of the LaRC/8-foot laminar-flow control (LFC) experiment was completed and the results are presented. The LFC airfoil used for this experiment was a swept, supercritical design which removed suction air through spanwise slots. The amplification of small disturbances by linear processes on a swept surface such as this can be due to either Tollmien-Schlichting (TS) and/or crossflow (CF) mechanisms. This study consists of the examination of these two instabilities by both the commonly used incompressible (SALLY and MARIA) analysis and the more involved compressible (COSAL) analysis. A wide range of experimental test conditions with variations in Mach number, Reynolds number, and suction distributions were available for this study. Experimentally determined transition locations were found from thin-film techniques and were used to correlate the n-factors at transition for the range of test cases.

The method of matched asymptotic expansions is used to analyze the interaction of a normal shock wave with an unseparated turbulent boundary layer on a flat surface at transonic speeds. The theory leads to a three-layer description of the interaction in the double limit of Reynolds number approaching infinity and Mach number approaching unity. The interaction involves an outer, inviscid rotational layer, a constant shear-stress wall layer, and a blending region between them.

A computer code is described which yields accurate solutions for a broad range of laminar, nonsimilar boundary layers, providing the inviscid flow field is known. The boundary layer may be subject to mass injection for perfect-gas, nonreacting flows. If no mass injection is present, the code can be used with either perfect-gas or real-gas thermodynamic models. Solutions, ranging from two-dimensional similarity solutions to solutions for the boundary layer on the Space Shuttle Orbiter during reentry conditions, have been obtained with the code. Comparisons of these solutions, and others, with solutions presented in the literature; and with solutions obtained from other codes, demonstrate the accuracy of the present code.

A controlled disturbance is generated in the freestream of the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) by focusing a high-powered Nd:YAG laser to create a laser-induced breakdown plasma. The plasma then cools, creating a freestream thermal disturbance that can be used to study receptivity. The freestream disturbance convects down-stream in the Mach-6 wind tunnel to interact with a flared cone model. The adverse pressure gradient created by the flare of the model is capable of generating second-mode instability waves that grow large and become nonlinear before experiencing natural transition in quiet flow. The freestream laser perturbation generates a wave packet in the boundary layer at the same frequency as the natural second mode, complicating time-independent analyses of the effect of the laser perturbation. The data show that the laser perturbation creates an instability wave packet that is larger than the natural waves on the sharp flared cone. The wave packet is still difficult to distinguish from the natural instabilities on the blunt flared cone.

The long-term goals are to understand and parameterize the basic physics of the wind, waves and associated temperature and humidity structure in the boundary layer over the ocean. This is done via direct measurements from the Research Platform FLIP of the wind profile, turbulence, waves and associated variables. Data from other experiments (coastal up-welling, tropical ocean (TOGA COARE), and over land) are used for comparison and extension to different regimes.

A new correlation of the roughness density effect on the intercept of the law of the wall and the equivalent sand roughness, has been developed. The data base used includes corrected results of Schlichting's rough channel experiment and those of twelve other studies. The roughness density parameter used is a product of the reference surface to the total frontal area ratio and a shape factor. For two dimensional roughness, the new parameter correlates data for circular wires with that of square bars. For three-dimensional roughnesses, it correlates results from all four types of roughness studied by Schlichting. The effect of three-dimensional roughness is a little smaller than that of two- dimensional roughness, having the same density parameter. A component build-up engineering analysis has been used to estimate the additional axial-force, due to grooves and threads, of projectiles. Comparison of this prediction technique and wind-tunnel data shows good agreement. Because the groove roughness, in the case considered here, is beyond the range of applicability of the new correlation, existing data are used to estimate an equivalent sand roughness. A turbulence model, based on that of Healzer and the equivalent sand roughness concept has been implemented in the Schiff and Steger PNS code. In order to numerically evaluate the conditions over a rough surface, a transition region between smooth and rough surfaces is introduced. This transition region allows the effect of the rough surface conditions to propagate into this viscous flow region in a realistic way. The predicted smooth body axial force coefficient, excluding the base drag, is five percent smaller than that measured in a wind tunnel test.

No abstract available

In a previous report the authors analyzed the stability of the lower branch solutions of the incompressible (M sub infinity = O) Falkner-Skan boundary layers. There a perturbation analysis to these boundary layers was performed resulting in the Rayleigh stability equation. Eigen value solutions were obtained for the Rayleigh equation for different adverse pressure gradient (beta) values. All retarded flows were found to be unstable for a small range of frequencies with the amplification factor increasing as the extent of reversed flow increased. In this report they have entended that work by including the effect of Mach number M sub infinity on the stability of adiabatic (S sub W = O) Falker-Skan equations for beta = -.04, -.08, -.12, -.16 and -.19884. We found out that in all these cases as the Mach number M sub infinity increases the instability of flow decreases. In most of the cases the instability almost completely disappeared at M sub infinity = 3.

The mechanism of sound generation and the kind of viscid interaction of vortices with airfoils in an airflow is investigated. Experiments have been performed in regular, quasi-stationary flow with vortices of a Karman vortex street. Depending on the dimensions of vortices and airfoils, their distance, and the flow Machnumbers, different kinds of upstream propagating sound waves occur. Keywords: Noise, Boundary layer, Vortex airfoil interaction, Secondary vortex, Vortex noise, Digital interferogram analysis, Fringe evaluation, Image processing.

A two-dimensional differential analysis is developed to approximate the turbulent boundary layer on a compressor blade element with strong adverse pressure gradients, including the separated region with reverse flow. The predicted turbulent boundary layer thicknesses and velocity profiles are in good agreement with experimental data for a cascade blade, even in the separated region.

A two-dimensional differential analysis is developed to approximate the turbulent boundary layer on a compressor blade element with strong adverse pressure gradients, including the separated region with reverse flow. The predicted turbulent boundary layer thicknesses and velocity profiles are in good agreement with experimental data for a cascade blade, even in the separated region.