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Aerodynamic noise research - acoustic environment due to separated flows and oscillating shocks and flow visualization experiments with supersonic separated turbulent flow
Comparisons are presented between numerical solutions of the Navier-Stokes and interacting boundary-layer equations for the subsonic separated turbulent flow over an axisymmetric boattail configuration for which experimental data was available. Comparisons were made using a standard algebraic turbulence model and its modification using a streamwise relaxation model. Grid size studies were performed to provide an approximate assessment of the sensitivity of each method to mesh size. In general, good agreement was observed between the solutions of the Navier-Stokes and interacting boundary-layer equations; however, the interaction approach responded more strongly to the modified turbulence model thereby giving better agreement with experimental data.
The first AIAA Drag Prediction Workshop, held in June 2001, evaluated the results from an extensive N-version test of a collection of Reynolds-Averaged Navier-Stokes CFD codes. The code-to-code scatter was more than an order of magnitude larger than desired for design and experimental validation of cruise conditions for a subsonic transport configuration. The second AIAA Drag Prediction Workshop, held in June 2003, emphasized the determination of installed pylon-nacelle drag increments and grid refinement studies. The code-to-code scatter was significantly reduced compared to the first DPW, but still larger than desired. However, grid refinement studies showed no significant improvement in code-to-code scatter with increasing grid refinement. The third Drag Prediction Workshop focused on the determination of installed side-of-body fairing drag increments and grid refinement studies for clean attached flow on wing alone configurations and for separated flow on the DLR-F6 subsonic transport model. This work evaluated the effect of grid refinement on the code-to-code scatter for the clean attached flow test cases and the separated flow test cases.
A method of controlling a shear layer for a fluid dynamic body introduces first periodic disturbances into the fluid medium at a first flow separation location. Simultaneously, second periodic disturbances are introduced into the fluid medium at a second flow separation location. A phase difference between the first and second periodic disturbances is adjusted to control flow separation of the shear layer as the fluid medium moves over the fluid dynamic body.
Report discusses numerical solutions of Navier-Stokes equations for viscous supersonic flow over fin protruding from flat plate. Two types of configurations considered: one with wedge-shaped fin (sharp edge facing upstream), other with blunt (rectangular parallelepiped) fin. Only steady flows considered. Computations performed to study phenomena involved in separations of three-dimensional flows.
A numerical algorithm that is second-order accurate in time has been developed for the conjugated problem of a separated, compressible flow field and a conductive solid body. The full two-dimensional time-dependent Navier-Stokes equations are coupled with the time-dependent energy equation for the solid body and are solved simultaneously. using implicit algorithms. The energy equation for the solid body may include arbitrarily distributed heat sources. The algorithm has been exmined for the case of two-dimensional supersonic compression-corner interaction, with a heat source embedded in the wall in the vicinity of the separation bubble and the attached boundary layer. The effect of the heat source on the flow field is studied for steady and transient cases.
In the near absence of gravity, a continuous-flow type of electrophoretic separator can be operated with a much thicker separation chamber than is possible under 1 g conditions. This should permit either better resolution or shorter separation time per unit of sample. An apparatus to perform experiments on sounding rockets is under development and will be described. The electrophoresis cell is 5 mm thick by 5 cm wide with 10 cm long electrodes. It is supplied with buffer, sample, and coolant at about 4 C through the use of a passive refrigerant system. UV sample detection and provision for recovery and cold storage of up to 50 sample fractions are now being added to the basic unit. A wide range of operating conditions are electronically programmable into the unit, even up to a short time before flight, and a further range of some parameters can be achieved by exchanging power supplies and by changing gears in the motor drive units of the pump. The preliminary results of some separation studies on various biological products using a commercially available electrophoretic separator are also presented.
Report discusses experiments in use of acoustic excitation to reduce separation of two-dimensional laminar flow about airfoils at low angles of attack. Purpose of investigation to gain better understanding of effect of excitation, scaling of parameters of excitation, and ranges of frequencies in which excitation reduces separation effectively.
A new method for predicting turbulent flow reattachment location has been developed. The method predicts the reattachment location using the surface static pressure distribution and requires only a single empirical parameter. The method is based on the well-known Stratford’s flow separation criterion, which results from an approximate solution to the equations of motion within the boundary layer. Because of the large pressure rise, the original Stratford separation criterion is not valid for reattaching flows. In order to apply the Stratford criterion for reattaching flows, the reference location was moved to the pressure drop conditions instead of pressure rise conditions. In addition, the pressure distribution was scaled using a technique that was originally introduced for supersonic base flows. A new empirical parameter was also obtained for the modified Stratford criterion. The method has been tested for 30 different separated flow configurations from the literature. The variety of test cases include subsonic flows, transonic flows, 2D flows, and axisymmetric flows, with experimental as well as high resolution numerical simulation results. The flow reattachment predictions with the new method agree very well with the reported data. The error is less than 5% and typically about 2%. The new method is an extremely simple and a quick way of finding the reattachment location. Given the complexities of the existing experimental techniques for finding the reattachment location, the method easily and quickly provides the flow reattachment location, which could be used to estimate the physical extent of separation bubbles, as a performance index for flow control methods, or to assess the predictive capability of turbulence models.
An investigation was conducted to determine the effect of the inlet Mach number and entrance-boundary-layer thickness on the performance of a 23 degree 21-inch conical-diffuser - tail-pipe combination with a 2:1 area ratio. The air flows used in this investigation covered an inlet Mach number range from 0.17 to 0.89 and corresponding Reynolds numbers of 1,700,000 to 7,070,000. Results are reported for two inlet-boundary-layer thicknesses. Over the entire range of flows, the mean value of the inlet displacement thickness is about 0.034 inch for the thinner inlet boundary layer and about 0.170 inch for the case of the thicker inlet boundary layer. The performance of the diffuser - tail-pipe combination is presented together with examples of longitudinal static-pressure distribution and the results of boundary-layer pressure surveys made at six points along the diffuser wall. The results indicated a progressive diminution of the static-pressure recovery and a steady increase in the total-pressure losses as the inlet Mach number was increased for both inlet-boundary-layer thicknesses. The ratio of actual static-pressure rise to that theoretically possible was much less and the total-pressure losses were greater for the case of the thicker inlet boundary layer throughout the speed range investigated. With the thinner inlet boundary layer, flow separation occurred at the diffuser exit at all inlet Mach numbers.Unseparated flow alternating with separated flow was observed near the inlet at the higher velocities. For the case of the thicker inlet boundary layer, the origin of the separated region occurred in the vicinity of the inlet-duct-diffuser junction section at all Mach numbers.
Viscous separated flow surrounding a hemisphere-cylinder body at angles of attack ranging up to 19 deg in transonic flow has been computed using an implicit, approximately-factored, partially flux-split algorithm. The resulting flowfield structures, including the vortical flow on the leeward side of the body and the three-dimensional separation patterns, have been investigated. The computed results show good qualitative and quantitative agreement with experimental data. Furthermore, visualization of the flowfield patterns has yielded insight into the behavior of the three-dimensional separated flow.
Subsonic wind tunnel tests were performed on solid and porous (22 percent) 5.0-caliber forebody models to assess the effect of free and fixed cross-flow separation on the effectiveness of passive porosity. The effectiveness of passive porosity to control the local pressure loading for forced cross-flow separation is found to be similar to that observed for the free cross-flow separation condition. It is also found that the effectiveness of passive porosity is significantly enhanced in the presence of large positive pressures on the porous surface.
Due to observations of unsteady flow in the Space Shuttle Main Engine fuel preburner injector element, several flow studies have been performed. Real time neutron radiography tests were recently completed. This technique provided real time images of MiL-c-7024 and Freon-22 flow through an aluminum liquid oxygen post model at three back pressures (0, 150, and 545 psig) and pressure drops up to 1000 psid. Separated flow appeared only while operating at back pressures of 0 and 150 psig. The behavior of separated flow was similar to that observed for water in a 3x acrylic model of the LOX post. On the average, separated flow appeared to reattach near the exit of the post when the ratio of pressure drop to supply pressure was about 0.75.
Current turbulence models, such as those employed in Reynolds-averaged Navier-Stokes CFD, are unable to reliably predict the onset and extent of the three-dimensional separated flow that typically occurs in wing-fuselage junctions. To critically assess, as well as to improve upon, existing turbulence models, experimental validation-quality flow-field data in the junction region is needed. In this report, we present an overview of experimental measurements on a wing-fuselage junction model that addresses this need. The experimental measurements were performed in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The model was a full-span wing-fuselage body that was configured with truncated DLR-F6 wings, both with and without leading-edge extensions at the wing root. The model was tested at a fixed chord Reynolds number of 2.4 million, and angles-of-attack ranging from -10 degrees to +10 degrees were considered. Flow-field measurements were performed with a pair of miniature laser Doppler velocimetry (LDV) probes that were housed inside the model and attached to three-axis traverse systems. One LDV probe was used to measure the separated flow field in the trailing-edge junction region. The other LDV probe was alternately used to measure the flow field in the leading-edge region of the wing and to measure the incoming fuselage boundary layer well upstream of the leading edge. Both LDV probes provided measurements from which all three mean velocity components, all six independent components of the Reynolds-stress tensor, and all ten independent components of the velocity triple products were calculated. In addition to the flow-field measurements, static and dynamic pressures were measured at selected locations on the wings and fuselage of the model, infrared imaging was used to characterize boundary-layer transition, oil-flow visualization was used to visualize the separated flow in the leading- and trailing-edge regions of the wing, and unsteady shear stress was measured at limited locations using capacitive shear-stress sensors. Sample results from the measurement techniques employed during the test are presented and discussed.
A technique for improving the numerical predictions of turbulent flows with the effect of streamline curvature is developed. Separated flows, the flow in a curved duct, and swirling flows are examples of flow fields where streamline curvature plays a dominant role. A comprehensive literature review on the effect of streamline curvature was conducted. New algebraic formulations for the eddy viscosity incorporating the kappa-epsilon turbulence model are proposed to account for various effects of streamline curvature. The loci of flow reversal of the separated flows over various backward-facing steps are employed to test the capability of the proposed turbulence model in capturing the effect of local curvature. The inclusion of the effect of longitudinal curvature in the proposed turbulence model is validated by predicting the distributions of the static pressure coefficients in an S-bend duct and in 180 degree turn-around ducts. The proposed turbulence model embedded with transverse curvature modification is substantiated by predicting the decay of the axial velocities in the confined swirling flows. The numerical predictions of different curvature effects by the proposed turbulence models are also reported.
A technique for improving the numerical predictions of turbulent flows with the effect of streamline curvature is developed. Separated flows and the flow in a curved duct are examples of flowfields where streamline curvature plays a dominant role. New algebraic formulations for the eddy viscosity incorporating the k-epsilon turbulence model are proposed to account for various effects of streamline curvature. The loci of flow reversal of the separated flows over various backward-facing steps are employed to test the capability of the proposed turbulence model in capturing the effect of local curvature.
This paper describes an apparatus for continuous flow electrophoresis (CFE), designed to separate macromolecules and cells at conditions of microgravity. In this CFE, buffer flows upward in a 120-cm long flow chamber, which is 16-cm wide x 3.0-mm thick in the microgravity version (and 6-cm wide x 1.5-mm thick in the unit-gravity laboratory version). Ovalbumin and rat serum albumin were separated in space (flight STS-4) with the same resolution of the two proteins achieved at 25 percent total w/v concentration that was obtained in the laboratory at 0.2 percent w/v concentration. Rat anterior pituitary cells, cultured human embryonic kidney cells, and canine Langerhans cells were separated into subpopulations (flight STS-8) more effectively than in unit gravity, with comparable resolution having been achieved at 100 times the concentration possible on earth.
Aerodynamic noise measurements have been obtained on three nose-cylinder configurations (15 deg cone-cylinder, biconic nose-cylinder and biconic nose-corrogated cylinder) in transonic flow. The noise environment at the foot of the terminal normal shock and its associated separated flow bubble were examined in detail. A peak in the frequency spectra of the separated flow region was observed which compares well with the predicted resonance frequency of the separated pocket using Trilling's incident shock model. As a consequence of the observed sound spectra and noise level, a flow model is postulated which consist of a cellular separated flow region. In the central portion of the cell the flow resembles the classical two-dimensional recirculating bubble, but each cell is vented by a pair of counter rotating vortices, which start at the surface and steam downstream.