Showing papers on "Freestream published in 1973"
TL;DR: In this article, an experimental investigation was made of the unsteady pressure on a circular cylinder in a free stream of constant and oscillating velocity, and it appeared that the independence of symmetric and antisymmetric modes in linear systems brings about this fortunate circumstance in these nonlinear flows.
Abstract: : An experimental investigation was made of the unsteady pressure on a circular cylinder in a free stream of constant and oscillating velocity. Consideration of possible resonant coupling mechanisms between the Karman vortex street fluctuations and freestream velocity oscillations showed such coupling to be insignificant, if any. It appears that the independence of symmetric and antisymmetric modes in linear systems brings about this fortunate circumstance in these nonlinear flows. Designs involving bluff bodies in flows with pulsations should therefore be much safer than previously believed. In such flows the Strouhal number evidently scales with the 'instantaneous' velocity when the pulsations are slow and with the mean velocity when the pulsating frequency is of the same order as the frequency of the vortex street. (Author)
21 citations
TL;DR: In this article, an analytical and experimental study of boundary layer flow over an aerodynamic surface rejecting heat to a cool environment was conducted following reentry of a Space Shuttle vehicle, showing that a surface to freestream temperature ratio greater than unity tended to destabilize the boundary layer, hastening transition and separation.
Abstract: Results of an analytical and experimental study of boundary layer flow over an aerodynamic surface rejecting heat to a cool environment. This occurs following reentry of a Space Shuttle vehicle. Analytical studies revealed that a surface to freestream temperature ratio, greater than unity tended to destabilize the boundary layer, hastening transition and separation. Therefore, heat transfer accentuated the effect of an adverse pressure gradient. Wind tunnel tests of a 0012-64 NACA airfoil showed that the stall angle was significantly reduced while drag tended to increase for freestream temperature ratios up to 2.2.
19 citations
TL;DR: The aerodynamic interference resulting from a jet issuing normal to the chordal plane of a twodimensional wing in a crossflow has been experimentally investigated in this article, where measurements of the interference surface pressure distribution on the wing and of the wing interference force and moment coefficients have been made for a systematic variation of jet exit location, jet exit diameter, wing angle-of-attack, and the ratio of the jet exit velocity to freestream velocity.
Abstract: The aerodynamic interference resulting from a jet issuing normal to the chordal plane of a twodimensional wing in a crossflow has been experimentally investigated. Measurements of the interference surface pressure distribution on the wing and of the wing interference force and moment coefficients have been made for a systematic variation of jet exit location, jet exit diameter, wing angle-of-attack, and the ratio of jet exit velocity to freestream velocity, A. A comparison of the contours of constant interference surface pressure on the wing lower surface with those for an infinite flat plate reveals that they are much the same for A > 6. The dissimilarity becomes greater as X is decreased, primarily through the growth of an extensive region of positive interference surface pressure forward of the jet on the wing. Interference lift losses of approximately the same magnitude for all geometries were observed for X > 6. However, a lift augmentation occurred for X < 6 which was attenuated by increases in angle-of-attack, forward movement of the jet exit location, and decreases in jet exit size. The data indicate that the character of the interference flow is distinctly different for high and low values of the velocity ratio.
17 citations
01 Jan 1973
TL;DR: In this article, a more accurate theoretical study of a solid polymeric fuel in a hot oxidizing flow field is carried out as an extension of the previous work reported in the Thirteenth Combustion Symposium.
Abstract: A more accurate theoretical study of ignition of a solid polymeric fuel in a hot oxidizing flow field is carried out as an extension of the previous work reported in the Thirteenth Combustion Symposium. In the present work, the complete non-similar calculation is carried out, eliminating the previously used local similarity approximation and the integral approximation to solve the condensed-phase energy equation. Comparison of the predictions based on the non-similar calculations indicates that both results predict similar trends of ignition behavior. However, the ignition delay time and the amount of downstream shift in the previous local similarity calculation differ significantly from those in the present non-similar calculation. This comparison has implications for other otheries of boundary layer combustion in the literature that utilize the self-similarity approximation. The flame-spreading behavior is also predicted as an extension of the ignition process. In agreement with the experimental observations reported in the previous work, results of the non-similar calculation predict that, in the case of high freestream oxygen concentration, the flame spreading downstream from the point of first ignition is slow. Reducing the oxygen concentration increases the flame-spreading speed in both directions (upstream and downstream). This aggreement between experiment and theory indicates that, when the flame spreads downstream, convective heat transfer is the dominant process in bringing the local surface element to vigorous pyrolysis, rather than conduction, and when the flame spreads upstream in the low freestream oxygen level case, the surface temperature ahead of the flame front is already high enough to cause spontaneous ignition, and the appearance of flame awaits the evolution of an abundant supply of fuel. Ignition and flame spreading are shown theoretically to depend sensitively on the thermal and transient properties of the diluent gas (e.g., helium, argon, nitrogen).
16 citations
TL;DR: In this article, the sharp leading edge problem for both monatomic and diatomic gases using the Boltzmann equation with the Bhatnagar-Gross-Krook type models as the governing equation and the discrete ordinate method with a closed-boundary value approach as a tool was studied.
Abstract: The sharp leading edge problem has been studied for both monatomic and diatomic gases using the Boltzmann equation with the Bhatnagar-Gross-Krook type models as the governing equation and the discrete ordinate method with a closed-boundary value approach as a tool. Plate length relative to the freestream mean free path is taken to be 52. The gas-surface interaction law is assumed to be diffuse reflection. The local distribution functions of molecular velocities and internal energies (for the diatomic gas) for the entire flowfield have been calculated for a freestream Mach number of 6.1. Comparisons are made between the calculated results and experimental data.
15 citations
14 citations
TL;DR: In this article, the chemically reacting boundary layer over a fuel plate was analyzed to show the effect of thermodynamic coupling at the solid-gas interface, where the fuel is assumed to undergo surface pyrolysis with subsequent reaction with the freestream oxidizer at a flame sheet.
Abstract: The chemically reacting boundary layer over a fuel plate was analyzed to show the effect of thermodynamic coupling at the solid-gas interface. The fuel is assumed to undergo surface pyrolysis with subsequent reaction with the freestream oxidizer at a flame sheet. Computed values of species and temperature profile, flame sheet and boundary-layer thickness, as well as surface blowing rates are presented for a variety of cases of stoichiometry, energy feedback, and pressure gradient. Although similarity approximations were utilized, the analysis compares favorably with results of regression studies during hybrid combustion.
14 citations
TL;DR: In this paper, the authors derived a correlation parameter for analysis of experimental Magnus force data, which is a function of angle of attack, the fineness ratio, and the Reynolds number for finless slender spinning circular cylinders.
Abstract: Time-dependent results of numerical finite-difference solution for the two-dimensional spinning circular cylinder are used to derive a correlation parameter for analysis of experimental Magnus force data. The correlation parameter, which is a function of angle of attack, the fineness ratio, and the Reynolds number, is derived on the basis of the impulse or cross-flow analogy. The correlation parameter, derived for finless slender spinning circular cylinders is shown to successfully correlate experimental data for fineness ratios from 6 to 24, for subsonic through supersonic freestream speeds, and for cross-flow Mach numbers up to 0.4 at transonic and supersonic speeds. Cy = Cyp = d = k = / = M p p t t/oo Uc V x 7 a v p Nomenclature two-dimensional lift coefficient, lift per unit span/(l/2)p Uc2 d local side-force coefficient, side-force per unit span/(l/2)p Uc2 d = Magnus force coefficient, 87/pl/^2 nd2p = body diameter = constant of proportionality = body length = freestream Mach number = spin rate = dimensionless spin rate, pd/211^ = time = freestream speed = component of freestream speed normal to body axis = tangential surface speed due to spin = distance along body axis from nose = Magnus force = angle of attack = kinematic viscosity = air density
14 citations
TL;DR: In this paper, an experiment is described which consists of monitoring the base pressure of an entry probe and deriving the freestream pressure profile of a planet through correlation curves which directly relate base pressure to free-stream pressure for the varying trajectory conditions.
Abstract: An experiment is described which consists of monitoring the base pressure of an entry probe and deriving the freestream pressure profile of a planet through correlation curves which directly relate base pressure to freestream pressure for the varying trajectory conditions. The experiment/technique is applicable for Mars, Venus, and Jupiter entry probes. The base pressure experiment offers distinct advantages to an entry probe mission such as a positive indication of where Mx = 1 occurs in the entry trajectory, a positive indication of boundary-layer transition onset, and a freestream pressure boundary condition at the A/x = 1.2 trajectory point independent of any other onboard or offboard measurements. The results of a recent slender cone R/V flight test have demonstrated the feasibility of the experiment by deriving the atmospheric pressure profile of Earth from base pressure measurements. Available flight and ground test base pressure data have been reviewed and an assessment made of which parameters are important to the base flow phenomena, which are well known and which require more investigaton to calibrate the experiment by obtaining additional data.
13 citations
TL;DR: In this article, a numerical study of the nature of the solution to the supersonic boundary-layer equations when the flow separates from the surface was made, and implicit finite difference solutions were obtained for linearly and quadratically retarded flows over flat plates for a Mach number range of 2 to 15.
Abstract: : A numerical study has been made of the nature of the solution to the supersonic boundary-layer equations when the flow separates from the surface. To this end, implicit finite difference solutions t the governing equations are obtained for linearly and quadratically retarded flows over flat plates for a Mach number range of 2 to 15. Both extremely hot and cold wall conditions are considered. Also, solutions for flow on a compression ramp are obtained with and without including the boundary layer's free interaction with the local freestream. The results of this study indicate that non-interacting flows are singular at separation and that free interaction effects reduces singularity. An alternate means of numerically smearing out the singularity and proceeding past separation is also presented.
10 citations
TL;DR: In this paper, the authors measured the flowfield structure and pressure and heat transfer distributions associated with small fins mounted near the base of a 7° half-angle cone, in a hypersonic freestream at a Mach number of 10, freeestream unit Reynolds numbers of 0.12 and 0.27.
Abstract: The flowfield structure (surface streamlines and shock patterns), and pressure and heat-transfer distributions associated with small fins mounted near the base of a 7° half-angle cone are measured. The tests are conducted in a hypersonic freestream at a Mach number of 10, freestream unit Reynolds numbers of 0.12 and 0.6 x 10 6/ foot, and wall to total temperature ratio of 0.27. Vehicle geometry (nose bluntness, and fin sweep, shape, thickness, and cant) and orientation (angles of attack and roll) are systematically varied. Fin leading-edge pressure and heat transfer agree with swept infinite cylinder predictions; fin side heating agrees with blunt swept slab theory. Flow patterns, heating and pressure in the fin-cone region are characteristic of corner flow geometries.
TL;DR: In this article, the authors present additional hypersonic boundary-layer transition measurements on a moderately cooled slender cone at a local Mach number considerably higher than previously reported in the literature.
Abstract: O the many factors influencing hypersonic boundary-layer transition, one of the most controversial is the effect of wall cooling (or heating). Some investigations" indicate that heat transfer has a moderate to strong effect on hypersonic boundarylayer transition. Conversely, other studies ~ indicate that heat transfer has no measurable effect. In addition, transition "reversal'' has been detected at extreme wall cooling for hypersonic,' as well as supersonic conditions. This Note presents additional hypersonic transition measurements on a moderately cooled slender cone at a local Mach number considerably higher than previously reported in the literature. The present results, combined with lower hypersonic Mach number data, allow speculation as to the influence of wall cooling on transition Reynolds number for the noise dominated hypersonic case. The tests were conducted in the Mach 20 leg of the Langley High Reynolds Number Helium Facility. This tunnel has an axisymmetric contoured nozzle with a 1.525-mdiam test section. Heat-transfer measurements were obtained at zero angle of attack on a 1.525-m-long, 2.87°-half-angle cone with a nose radius of 0.010 cm. The model was instrumented with thermocouples for determining heating rates (from which the transition location was determined). Local unit Reynolds number for the tests was essentially constant at about-Re/m& 37 x 10. The helium freestream flow was unheated with Tt ^ & 300°K. Previous transition data for the same model and facility combination, for the uncooled case, are available in Kef. 9. Cooling of the model surface was accomplished by venting cold helium vapor into the interior of the cone. The cold helium vapor entered through a tube which ran from the helium supply through the base of the model, and terminated internally about midway along the cone length. The helium vapor exited at this point and distributed itself inside the cone, eventually passing into
TL;DR: The sensitivity of boundary-layer transition on hypersonic lifting geometries at high angles of attack (a) to relatively small amounts of surface roughness (irregularities) is illustrated in this paper.
Abstract: T purpose of this note is to illustrate the sensitivity of boundary-layer transition on hypersonic lifting geometries at high angles of attack (a) to relatively small amounts of surface roughness (irregularities). Tests were conducted in the AEDC-VKF Hypervelocity Wind Tunnel F (Hotshot), on a 0.011-scale model of the McDonnell Douglas (MDAC) Phase B, STS Orbiter configuration at simulated hypersonic re-entry conditions. Measurements of laminar, transitional, and turbulent flow heating rate distributions and the location of boundary-layer transition were obtained. Selected data are presented for a freestream Mach number (MM) of 10.8 and a-40°.
01 Nov 1973
TL;DR: In this paper, the use of film cooling for protecting a nose cone in a hypersonic Mach 6 freestream was investigated experimentally in a Mach 6 wind tunnel with a contoured axisymmetric nozzle.
Abstract: : The use of film cooling for protecting a nose cone in a hypersonic Mach 6 freestream was investigated experimentally. Tests were performed in a Mach 6 wind tunnel with a contoured axisymmetric nozzle. Downstream and upstream tangential slot injections were applied to investigate the film cooling effectiveness on the surface of a nose cone. Multiple tangential downstream slot injections were used to cool the surface of the blunt nose cone while tangential upstream slot injections were used for the surface cooling of a sharp nose cone. Air at a stagnation temperature of 530 degrees R was used as an injectant. Surface distributions of the heat transfer rates and static pressures were measured for different injection mass flow rates.
TL;DR: In this paper, a 10% scale model of the Viking Lander with freestream velocity conditions from 37 to 75 ft/sec was used for simulation and the simulation test data were close to the theoretical predictions of Hill et al.
Abstract: Simulation experiments were performed in the NASA Langley Full Scale Tunnel to investigate the retroplume heating rates encountered by a lander vehicle during a terminal descent to a planetary surface. A 10% scale model of the Viking Lander with freestream velocity conditions from 37 to 75 ft/sec was used for simulation. The simulation test data were close to the theoretical predictions of Hill et al. (1971).
01 Feb 1973
TL;DR: In this paper, a two-dimensional theory for the unsteady flow disturbances caused by aeroelastic deformations of a thick wing at high subsonic freestream Mach numbers, having a single, locally supersonic steady flow region adjacent to the low pressure side of the wing, is considered.
Abstract: A two-dimensional theory is considered for the unsteady flow disturbances caused by aeroelastic deformations of a thick wing at high subsonic freestream Mach numbers, having a single, internally embedded supercritical (locally supersonic) steady flow region adjacent to the low pressure side of the wing. The theory develops a matrix of unsteady aerodynamic influence coefficients (AICs) suitable as a strip theory for aeroelastic analysis of large aspect ratio thick wings of moderate sweep, typical of a wide class of current and future aircraft. The theory derives the linearized unsteady flow solutions separately for both the subcritical and supercritical regions. These solutions are coupled together to give the requisite (wing pressure-downwash) AICs by the intermediate step of defining flow disturbances on the sonic line, and at the shock wave; these intermediate quantities are then algebraically eliminated by expressing them in terms of the wing surface downwash.
01 Nov 1973
TL;DR: In this article, the authors investigated several interaction theories for high speed vehicles and found that the deceleration of the fluid by the viscous processes in a boundary layer generally produces very high temperatures.
Abstract: : At hypersonic speeds, the deceleration of the fluid by the viscous processes in a boundary layer generally produces very high temperatures. One result of these high temperatures is an increase in the thickness encountered at the same freestream Reynolds number at lower speeds. Viscous interaction is one of a number of important high speed phenomena that require proper evaluation prior to making any realistic aerodynamic performance estimate for a high speed vehicle. It is the purpose of the report to experimentally investigate several interaction theories.
01 Jul 1973
TL;DR: In this paper, an implicit finite-difference method has been developed for the solution of the compressible boundary-layer equations, which is applied to tangential slot injection into supersonic turbulent boundary layer flows.
Abstract: An implicit finite-difference method has been developed for the solution of the compressible boundary-layer equations. This method is applied to tangential slot injection into supersonic turbulent boundary-layer flows. In addition, the effects induced by the interaction between the boundary-layer displacement thickness and the external pressure field are considered. Three different eddy viscosity models have been used to specify the turbulent momentum exchange. One model depends on the species concentration profile, and the species conservation equation has been included in the system of governing partial differential equations. For air injected into air, the freestream and injected gases are treated as separate species which have common fluid properties. Calculations were made and results were compared with experimental data at freestream Mach numbers of 2.4 and 6.0 and with results of another finite-difference method. Good agreement was obtained for the reduction of wall skin friction with slot injection.
01 Aug 1973
TL;DR: In this article, an experimental program was conducted to evaluate the use of encapsulated liquid crystals as a means of determining temperature profiles in regions of interfering flows on wind tunnel models at supersonic speeds.
Abstract: : An experimental program was conducted to evaluate the use of encapsulated liquid crystals as a means of determining temperature profiles in regions of interfering flows on wind tunnel models at supersonic speeds. The tests were conducted on both plane surfaces and on an ogive cylinder model equipped with 3 dimensional shock generators at Mach numbers of 1.89 and 3.00 in the Trisonic Gasdynamic Facility of the Air Force Flight Dynamics Laboratory. Freestream Reynolds number were 1.0 and 3.0 million per foot. Results show that models coated with encapsulated liquid slurry and overcoated with a protective plastic film provided better temperature profiles than models equipped with paper substrated liquid crystals. Temperature profiles were easily discernible over the entire surface of the ogive cylinder nodal within viewing range of the camera, and specular reflections on the non-planer surfaces did not invalidate any data.