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Free laminar boundary layer flow between parallel streams of different magnetic fields and temperatures for incompressible viscous fluid
The Transonic Analysis of a Wing and Fuselage with Interacted Viscous Effects (TAWFIVE) was developed. A finite volume full potential method is used to model the outer inviscid flow field. First-order viscous effects are modeled by a three dimensional integral boundary layer method. Both turbulent and laminar boundary layers are treated. Wake thickness and curvature effects are modeled using a two dimensional strip method. A very brief discussion of the engineering aspects of the program is given. The input and use of the program are covered in great detail.
A program for the Transonic Analysis of a Wing and Fuselage with Interacted Viscous Effects (TAWFIVE) was improved by the incorporation of multigrid and a method to specify lift coefficient rather than angle-of-attack. A finite volume full potential multigrid method is used to model the outer inviscid flow field. First order viscous effects are modeled by a 3-D integral boundary layer method. Both turbulent and laminar boundary layers are treated. Wake thickness effects are modeled using a 2-D strip method. A brief discussion of the engineering aspects of the program is given. The input, output, and use of the program are covered in detail. Sample results are given showing the effects of boundary layer corrections and the capability of the lift specification method.
Investigation of the effects of surface temperature on the stability and transition of laminar boundary layers
Flat plate and stagnation flow heat transfer coefficients, similarity solutions of the laminar boundary layer for air in dissociation equilibrium and calculation of laminar heat-transfer distribution on blunt three-dimensional bodies in high speed flow
A simulation of steady transonic viscous-inviscid interactions through a combination of zonal solution methods (involving different equation sets and numerical regions for various flow regions) is used in an investigation of a fluid mechanics problem in which the pressure distribution is determined by the interaction between transonic inviscid flow and an inner laminar viscous layer. The results obtained are in good agreement with laminar experimental data. Both attached and separated boundary layer flows are considered for either weak or strong interactions.
As an alternative to pure finite-difference techniques, the Method of Integral Relations (MIR) is formulated for three-dimensional boundary-layer flows with separation, and applied to two problems. The first concerns an incompressible laminar flow over a plate with an attached cylinder, where the potential solution of the flow round the circular cylinder is used as the boundary condition. In the second application the solution to the three-dimensional boundary-layer flow over blunt bodies is discussed and results for an ellipsoid of revolution at 30 deg incidence are presented. In both cases the results are found to be in satisfactory agreement with those obtained by finite-difference methods.
Noise produced by cold elliptic jets was studied experimentally using a 4 cm air jet in an anechoic chamber by varying the aspect ratio, nozzle exit boundary layer, and Mach number. Elliptic geometry is found to accentuate the sound produced by pairing of the shear layer vortices in the laminar boundary jet. Calculation of the pressure field produced by the pairing of elliptic vortex rings shows that the ratio of sound intensities produced by the pairing of inviscid elliptic vortex filament rings should be equal to their aspect ratio, the minor axis producing greater sound. Tripping the nozzle boundary layer changes the instability of the shear layer and removes the laminar vortices. If the sound produced by pairing and/or breakdown is considered to be the difference between laminar and turbulent jet noise, the pairing mechanism is seen to be very significant in laminar boundary layer jets.
A technique is developed to measure the local convective heat transfer coefficient on a model surface in a supersonic flow field. The technique uses a laser to apply a discrete local heat flux at the model test surface, and an infrared camera system determines the local temperature distribution due to heating. From this temperature distribution and an analysis of the heating process, a local convective heat transfer coefficient is determined. The technique was used to measure the load surface convective heat transfer coefficient distribution on a flat plate at nominal Mach numbers of 2.5, 3.0, 3.5, and 4.0. The flat plate boundary layer initially was laminar and became transitional in the measurement region. The experimental results agreed reasonably well with theoretical predictions of convective heat transfer of flat plate laminar boundary layers. The results indicate that this non-intrusive optical measurement technique has the potential to obtain high quality surface convective heat transfer measurements in high speed flowfields.
A technique is developed to measure the local convective heat transfer coefficient on a model surface in a supersonic flow field. The technique uses a laser to apply a discrete local heat flux at the model test surface, and an infrared camera system determines the local temperature distribution due to the heating. From this temperature distribution and an analysis of the heating process, a local convective heat transfer coefficient is determined. The technique was used to measure the local surface convective heat transfer coefficient distribution on a flat plate at nominal Mach numbers of 2.5, 3.0, 3.5, and 4.0. The flat plate boundary layer initially was laminar and became transitional in the measurement region. The experimentally determined convective heat transfer coefficients were generally higher than the theoretical predictions for flat plate laminar boundary layers. However, the results indicate that this nonintrusive optical measurement technique has the potential to measure surface convective heat transfer coefficients in high speed flow fields.
A technique is developed to measure the local convective heat transfer coefficient on a model surface in a supersonic flow field. The technique uses a laser to apply a discrete local heat flux at the model test surface, and an infrared camera system determines the local temperature distribution due to the heating. From this temperature distribution and an analysis of the heating process, a local convective heat transfer coefficient is determined. The technique was used to measure the local surface convective heat transfer coefficient distribution on a flat plate at nominal Mach numbers of 2.5, 3.0, 3.5, and 4.0. The flat plate boundary layer initially was laminar and became transitional in the measurement region. The experimentally determined convective heat transfer coefficients were generally higher than the theoretical predictions for flat plate laminar boundary layers. However, the results indicate that this nonintrusive optical measurement technique has the potential to measure surface convective heat transfer coefficients in high-speed flowfields.
A computer program was developed to do the calculations for two-dimensional or axisymmetric configurations from low speeds to hypersonic speeds with arbitrary streamwise pressure, temperature, and Mach number distributions. Options are provided for obtaining initial conditions either from experimental information or from a theoretical similarity solution. The transition region can be described either by an arbitrary distribution of intermittency or by a function based on Emmons' probability theory. Correlations were developed for use in estimating the parameters of the theoretical intermittency function. Correlations obtained from other sources are used for estimating the transition point. Comparisons were made between calculated and measured boundary layer quantities for laminar, transitional, and turbulent flows on flat plates, cones, cone flares, and a waisted body of revolution. Excellent agreement was obtained between the present theory and two other theories based on the method of finite differences. The intermittency required to reproduce some experimental heat transfer results in hypersonic flow was found to be quite different from the theoretical function. It is suggested that the simple probability theory of Emmons may not be valid for representing the intermittency of hypersonic transitional boundary layers and that the program could be useful as a tool for detailed study of the intermittency of the transition region.
An investigation of the three important factors that determine convective heat-transfer characteristics at supersonic speeds, location boundary-layer transition, recovery factor, and heat-transfer parameter has been performed at Mach numbers from 1.49 to 1.18. The bodies of revolution that were tested had, in most cases, laminar boundary layers, and the test results have been compared with available theory. Boundary-layer transition was found to be affected by heat transfer. Adding heat to a laminar boundary layer caused transition to move forward on the test body, while removing heat caused transition to move rearward. These experimental results and the implications of boundary-layer-stability theory are in qualitative agreement.
An investigation of the interaction of a laminar boundary layer and surface static-pressure orifices was conducted on a natural laminar flow airfoil. The effects of both favorable and adverse pressure gradients were studied at Reynolds numbers based on airfoil chord from approximately 500,000 to 6,000,000 with Mach number varying accordingly from about 0.03 to 0.42. The smaller the diameter of the orifice, the less likely it will cause premature transition. Locating the orifices in a chord-wise row aligned with the flow appears to have an additive, adverse effect on transition. Tube length-to-orifice diameter ration does not seem to have a consistent influence on transition.
Three-dimensional laminar boundary layer on body of revolution at small incidence at its stability
Instability and transition to turbulence in separated shear layer produced by laminar boundary layer separation from rearward-facing step, noting periodic spanwise structure
Theoretical results are presented for the structure of the hypersonic flow field of a blunt slab delta wing at moderately high angle of attack. Special attention is devoted to the interaction between the boundary layer and the inviscid entropy layer. The results are compared with experimental data. The three-dimensional inviscid flow is computed numerically by a marching finite difference method. Attention is concentrated on the windward side of the delta wing, where detailed comparisons are made with the data for shock shape and surface pressure distributions. Surface streamlines are generated, and used in the boundary layer analysis. The three-dimensional laminar boundary layer is computed numerically using a specially-developed technique based on small cross-flow in streamline coordinates. In the rear sections of the wing the boundary layer decreases drastically in the spanwise direction, so that it is still submerged in the entropy layer at the centerline, but surpasses it near the leading edge. Predicted heat transfer distributions are compared with experimental data.
Key data from a comprehensive airfoil broadband self-noise study are reported. Attention here is restricted to two-dimensional sharp trailing-edge models. The models include seven NACA 0012 airfoil sections and five 'flat plate' sections with chordlengths ranging from 2.54 to 60.96 cm. Testing parameters include flow velocity, angle of attack to the flow, and boundary layer turbulence through natural transition and through tripping. Detailed aerodynamic measurements, of pertinence to the scaling problem of airfoil self-noise, were conducted in the near-wake of the trailing edges. Presented are mean and rms turbulent velocity profiles as well as boundary layer thicknesses and integral thickness parameters for a large range of conditions. The noise spectra of the self-noise sources were determined by the use of a cross-spectral technique. The spectra were normalized using the measured aerodynamic parameters in order to evaluate the most common scaling law now in use. An examination of the Reynolds number dependence of the overall self-noise levels has revealed a new and useful scaling result. This result appears to quantify the transition between turbulent-boundary-layer trailing edge noise and laminar-boundary-layer vortex shedding noise.