The compressible turbulent shear layer: an experimental study

The growth rate and turbulent structure of the compressible, plane shear layer are investigated experimentally in a novel facility. In this facility, it is possible to flow similar or dissimilar gases of different densities and to select different Mach numbers for each stream. Ten combinations of gases and Mach numbers are studied in which the free-stream Mach numbers range from 0.2 to 4. Schlieren photography of 20-ns exposure time reveals very low spreading rates and large-scale structures. The growth of the turbulent region is defined by means of Pitot-pressure profiles measured at several streamwise locations. A compressibility-effect parameter is defined that correlates and unifies the experimental results. It is the Mach number in a coordinate system convecting with the velocity of the dominant waves and structures of the shear layer, called here the convective Mach number. It happens to have nearly the same value for each stream. In the current experiments, it ranges from 0 to 1.9. The correlations of the growth rate with convective Mach number fall approximately onto one curve when the growth rate is normalized by its incompressible value at the same velocity and density ratios. The normalized growth rate, which is unity for incompressible flow, decreases rapidly with increasing convective Mach number, reaching an asymptotic value of about 0.2 for supersonic convective Mach numbers

MODELING AND SIMULATION OF MIXING LAYER FLOWS FOR ROCKET ENGINE FILM COOLING

Film cooling has been selected for the thermal protection of the composite nozzle extension of the J-2X engine which is currently being developed for the second stage of NASA's next generation launch vehicle, the Ares I rocket. However, several challenges remain in order to achieve effective film cooling of the nozzle extension and to ensure its safe operation. The extreme complexity of the flow (three-dimensional wakes, lateral flows, vorticity, and flow separation) makes predicting film cooling performance difficult. There is also a dearth of useful supersonic film cooling data available for engineers to use in engine design and a lack of maturity of CFD tools to quantitatively match supersonic film cooling data. This dissertation advances the state of the art in film cooling by presenting semi-empirical analytical models which improve the basic physical understanding and prediction of the effects of pressure gradients, compressibility and density gradients on film cooling effectiveness. These models are shown to correlate most experimental data well and to resolve several conflicts in the open literature. The core-to-coolant stream velocity ratio, R, and the Kays acceleration parameter, KP, are identified as the critical parameters needed to understand how pressure gradients influence film cooling performance. The convective Mach number, Mc, the total temperature ratio, Ω0, and the Mach number of the high speed stream, MHS, are shown to be important when explaining the effects of compressibility and density gradient on film cooling effectiveness. An advance in the simulation of film cooling flows is also presented through the development of a computationally inexpensive RANS methodology capable of correctly predicting film cooling performance under turbulent, subsonic conditions. The subsonic simulation results suggest that it in order to obtain accurate predictions using RANS it is essential to thoroughly characterize the turbulent states at the inlet of the coolant and core streams of the film cooling flow. The limitations of this approach are established using a Grid Convergence Index (GCI) Test and a demonstration of the extension of this RANS methodology to supersonic conditions is presented

Institute for Computational Mechanics in Propulsion (ICOMP)

The Institute for Computational Mechanics in Propulsion (ICOMP) is a combined activity of Case Western Reserve University, Ohio Aerospace Institute (OAI) and NASA Lewis. The purpose of ICOMP is to develop techniques to improve problem solving capabilities in all aspects of computational mechanics related to propulsion. The activities at ICOMP during 1991 are described

Performance and boundary-layer evaluation of a sonic inlet

Tests were conducted to determine the boundary layer characteristics and aerodynamic performance of a radial vane sonic inlet with a length/diameter ratio of 1 for several vane configurations. The sonic inlet was designed with a slight wavy wall type of diffuser geometry, which permits operation at high inlet Mach numbers (sufficiently high for good noise suppression) without boundary layer flow separation and with good total pressure recovery. A new method for evaluating the turbulent boundary layer was developed to separate the boundary layer from the inviscid core flow, which is characterized by a total pressure variation from hub to tip, and to determine the experimental boundary layer parameters

Research in Natural Laminar Flow and Laminar-Flow Control, part 2

Part 2 of the Symposium proceedings includes papers addressing various topics in basic wind tunnel research/techniques and computational transitional research. Specific topics include: advanced measurement techniques; laminar flow control; Tollmien-Schlichting wave characteristics; boundary layer transition; flow visualization; wind tunnel tests; flight tests; boundary layer equations; swept wings; and skin friction

Simulation And Modeling Of Hypersonic Turbulent Boundary Layers Subject To Adverse Pressure Gradients Due To Concave Streamline Curvature

Direct numerical simulations (DNS) of adverse-pressure-gradient turbulent boundary layers over a planar concave wall are presented for a nominal freestream Mach number of 5, with the objective of assessing the limitations of the currently available Reynolds-averaged Navier-Stokes (RANS) models. The wall geometry and flow conditions of the DNS are representative of the experimental data for a Mach 4.9 turbulent boundary layer that was tested on a two-dimensional planar concave wall model in the high-speed blow-down wind tunnel located at the National Aerothermochemistry Laboratory at Texas A&M University (TAMU). The DNS was validated against the experimental results of TAMU for the same flow conditions and wall geometry. An analysis of the DNS datasets was also conducted to provide an assessment of the validity of Morkovin\u27s hypothesis and the strong Reynolds analog for turbulence subject to mechanical nonequilibrium. In addition to the DNS results, RANS predictions are obtained by using the Baldwin-Lomax (BL), Spalart-Allmaras (SA), and thek-ω SST turbulence models. The comparisons between RANS and DNS showed little impact of an adverse pressure gradient on the accuracy of these models, at least up to an incompressible Clauser pressure gradient parameter of ßinc=1•22 While the Boussinesq assumption provided reasonable predictions for the Reynolds shear stress, it failed to adequately predict the normal components of the Reynolds stress

Critical review of the trailing edge condition in steady and unsteady flow. Blade flutter in compressors and fans: Numerical simulation of the aerodynamic loading

Existing interpretations of the trailing edge condition, addressing both theoretical and experimental works in steady, as well as unsteady flows are critically reviewed. The work of Kutta and Joukowski on the trailing edge condition in steady flow is reviewed. It is shown that for most practical airfoils and blades (as in the case of most turbomachine blades), this condition is violated due to rounded trailing edges and high frequency effects, the flow dynamics in the trailing edge region being dominated by viscous forces; therefore, any meaningful modelling must include viscous effects. The question of to what extent the trailing edge condition affects acoustic radiation from the edge is raised; it is found that violation of the trailing edge condition leads to significant sound diffraction at the tailing edge, which is related to the problem of noise generation. Finally, various trailing edge conditions in unsteady flow are discussed, with emphasis on high reduced frequencies

Direct numerical simulation of pressure fluctuations induced by supersonic turbulent boundary layers

Direct Numerical Simulations are used to generate a database of high-speed zero-pressure-gradient turbulent boundary layers developing spatially over a flat plate with nominal freestream Mach number ranging from 2:5 to 14 and wall-to-recovery temperature ranging from 0:18 to 1:0. The flow conditions of the DNS are representative of the operational conditions of the Purdue Mach 6 quiet tunnel, the Sandia Hypersonic Wind Tunnel at Mach 8, and the AEDC Hypervelocity Tunnel No. 9 at Mach 14. The DNS database is used to gauge the performance of compressibility transformations, including the classical Morkovin\u27s scaling and strong Reynolds analogy as well as the newly proposed mean velocity and temperature scalings that explicitly account for wall heat flux, examine the pressure fluctuations generated by the turbulent boundary layers. The unsteady pressure field is analyzed at multiple wall-normal locations, including those at the wall, within the boundary layer (including inner layer, the log layer, and the outer), and in the free stream. The statistical and structural variations of pressure fluctuations as a function of wall-normal distance are highlighted. The simulations show that the dominant frequency of boundary-layer-induced pressure fluctuations shifts to lower frequencies as the location of interest moves away from the wall. The pressure structures within the boundary layer and in the free stream evolve less rapidly as the wall temperature decreases, resulting in an increase in the decorrelation length of coherent pressure structures for the colder wall case. The pressure structures propagate with similar speeds for both wall temperatures. Acoustic sources are largely concentrated in the near-wall region; wall cooling most significantly influences the nonlinear (slow) component of the acoustic source term by enhancing dilatational fluctuations in the viscous sublayer while damping vortical fluctuations in the buffer and log layers. Precomputed flow statistics, including Reynolds stresses and their budgets, are available at the website of the NASA Langley Turbulence Modeling Resource --Abstract, page iv