Search NASA⌕ Search

SEARCH · Search NASA

Results for “FLOW, SUPERSONIC”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 415 records · Page 23
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Viscous calculations of supersonic flows past cylinders at angles of attack

Two- and three-dimensional separated flows around cylindrical sections and a body-of-revolution at angles of attack from -10 deg to +20 deg are computed. The blunt-nose-cylinders chosen for the analyses have different configurations, and one of them has a boat-tail. The freestream Mach numbers are 1.5 and 1.6, the unit Reynolds number and the total temperature are 2 million/ft, and 585 R, respectively. Among the possible areas of application for this study are the aerodynamics of separating stores and missiles. The Reynolds-averaged Navier-Stokes equations are solved by an approximately-factored, implicit, upwind-biased, finite-volume algorithm. The convergence to steady state was accelerated using a multigrid strategy. The results helped understanding the structure of the flowfield with varying incidence and configuration, and the suitability of the computational algorithm used for this class of flows. A two-layer algebraic eddy viscosity model is used for the two-dimensional turbulent flows. Laminar three-dimensional calculations are benchmarked through comparisons with wind tunnel test results.

Baysal, O.↗
NASA NTRS1988

Computations of vortical supersonic flows over a blunt-nose-cylinder at high angles of attack

Three-dimensional, viscous, and separated flows over a blunt-nose-cylinder at 20, 32, and 44-deg attack angles were computed. The approaching freestream was at a Mach number of 1.6 and a unit Reynolds number of 2 x 10 to the 6th/ft with a total temperature of 585 R. The cylinder used for the computations had a length-to-diameter ratio of 6.67 with a base diameter of 3 in. The flowfield was dominated by large-scale and multiple vortices generated by crossflow separation. The effect of turbulence on the flow structure of one case was modeled algebraically with modifications to correct the length and velocity scales in the regions of separation. The mass averaged Navier-Stokes equations were solved by an approximately factored, upwind-biased, implicit, finite volume scheme. The initialization of the flows was enhanced by a mesh sequencing strategy applied to the diagonalized form of the discretized equations. The convergence to steady-state was accelerated by a multigrid algorithm and using the block inversions for the discretized equations. Calculations were compared with experimental results.

Baysal, O.↗
NASA NTRS1989

Supersonic flow computations by two-equation turbulence modeling

In the present work a solver for the Reynolds averaged compressible Navier-Stokes equations, to compute high speed turbulent flows characterized by interacting shock waves and viscous layers, is presented. A k-epsilon turbulence model that accounts for compressibility effects is developed. The numerical algorithm is based on a finite volume multistage Runge Kutta technique that is explicit for the solution of the mean flow variables, and implicit for the solution of the k-epsilon equations. The model is validated by extensive comparison with experimental results of flows over compression ramps characterized by interacting shock waves/boundary layers.

Grasso, Francesco↗
NASA NTRS1989

Low angle injection into a supersonic flow

A study of helium jets injected into a Mach 3 airflow through circular sonic nozzles angled downstream at 15 and 30 deg is presented. An aspirating hot-film probe was employed to measure helium concentrations. The results of Schlieren and nonoshadowgraph photography show the presence of large-scale turbulent structures in the high pressure jets. Concentration data reveal some asymmetries both in the cross section of the jet and in the trajectory of the jet relative to the centerline of the jet. Plotting the decay of maximum concentration with axial distance, it is found that the mixing rate can be increased significantly by increasing the pressure ratio. These low angle jets display good cross-stream penetration.

Mays, R. B.↗
NASA NTRS1990

Transient behavior of supersonic flow through inlets

A solution technique to compute inlet flow behavior is presented. The phenomena of inlet unstart and restart are investigated using a flux-split procedure applied to the Euler and Reduced Navier Stokes (RNS) equations. A time consistent direct sparse matrix solver in conjunction with a domain decomposition strategy is applied to compute the transient flow behavior both internal and external to the inlet. Time varying shocks and time varying recirculation regions are efficiently analyzed. The code is quite general and is suitable for the computation of flow for a wide variety of geometries and over a wide range of Mach and Reynolds numbers.

Pordal, H. S.↗
NASA NTRS1991

Effects of yaw on low angle injection into a supersonic flow

This paper presents the results of a study of transverse gas injection into supersonic/hypersonic streams with low downstream transverse angles in addition to yaw angles varying from zero to approximately 30 deg. The primary data are concentration measurements, with nanoshadowgraphs and oil flow visualization pictures also included. Two sets of experiments were performed. The first set studied the effects of yaw angle, specifically beta = 15 and 28 deg, on a helium injector with a 30-deg transverse angle in a Mach 3 freestream. Axial measurement stations were x/d = 30, 50, and 100. It was found that, as beta was increased, the maximum concentration mixing rate did not vary, but the jet core penetration decreased more at beta = 15 deg than at beta = 28 deg. A shearing effect between the portion of the jet in the boundary layer and the portion in the freestream increased the area of a typical constant concentration contour. The second set of experiments, conducted at NASA Langley, studied the effect of yawed injection at a transverse angle of 15 deg in a Mach 6 flow. Axial stations of x/d = 20, 40, 60, and 80 were used. A yaw angle of beta = 15 deg was found to decrease both the jet core mixing rate and penetration. The primary benefit of yaw was to increase lateral spreading. For similar injection conditions, the results show less near-field mixing at Mach 6 than Mach 3, but a faster mixing rate in the far-field at Mach 6.

Thomas, R. H.↗
NASA NTRS1991

Continuous wave laser absorption techniques for gasdynamic measurements in supersonic flows

Line-of-sight measurements of velocity, temperature, pressure, density, and mass flux were performed in a transient shock tube flow using three laser absorption schemes. All methods employed an intracavity-doubled ring dye laser tuned to an OH transition at 306 nm. In the first scheme, the gas was labeled by 193.3-nm excimer photolysis of H2O, and the passage of the generated OH was detected downstream. In the second method, the laser was tuned at a rate of 3 kHz over the R1(7) and R1(11) line pair, and absorption was simultaneously monitored at 90 and 60 deg with respect to the flow. Velocity was determined from the Doppler shift of the profiles and the temperature from the intensity ratio of the lines. Pressure was determined from both the magnitude of absorption and the collisional broadening. In the third method, the laser wavelength was fixed at a single frequency, and a continuous measurement of velocity and pressure was obtained using the signals from the two beam paths. All methods gave results which compare favorably to calculated values.

Davidson, David F.↗
NASA NTRS1991

LV software support for supersonic flow analysis

During 1991, the software developed allowed an operator to configure and checkout the TSI, Inc. laser velocimeter (LV) system prior to a run. This setup procedure established the operating conditions for the TSI MI-990 multichannel interface and the RMR-1989 rotating machinery resolver. In addition to initializing the instruments, the software package provides a means of specifying LV calibration constants, controlling the sampling process, and identifying the test parameters.

Bell, William A.↗
NASA NTRS1992

Finite element nonlinear panel flutter with arbitrary temperatures in supersonic flow

A finite element frequency domain method for predicting nonlinear flutter response of panels with temperature effects is presented. By using the principle of virtual work, the element nonlinear stiffness formulation for a panel under a combined thermal and aerodynamic loads is derived on the bases of von Karman's large deflection plate theory, the first-order piston theory aerodynamics and the quasi-steady thermal stress theory. The system equations of motion can be mathematically separated into two sets of equations and then solved in sequence. The first set of equations yields the panel thermal-aerodynamic equilibrium and the second set of equations of motion leads to the flutter limit-cycle oscillations. Stability and flutter boundaries can also be obtained from the two sets of system equations. Finite element large amplitude limit-cycle flutter results at different uniform temperatures are obtained for a simply supported square panel and are compared with existing Galerkin/time integration and other finite element solutions. Effects of nonuniform temperature distributions, panel length-to-width ratios, and boundary conditions on flutter responses of rectangular and triangular panels are presented.

Xue, David Y.↗
NASA NTRS1991

Computational methods for shock waves in three-dimensional supersonic flow

A central artificial-viscosity and an upwind-biased difference method are contrived to solve the Euler equations for flowfields over typical spacecrafts. The spatial discretization is based on either nodal or cell-vertex formulation in the domain extending from free stream to the end of the vehicle. The outer boundary is treated as a bow shock in the first method but is placed in the free stream in the second, which captures both bow and internal shocks using an approximate Riemann solver based on high-order extrapolation to the cell face. These methods were tested for the Shuttle and Hermes orbiters at wind-tunnel conditions and angles of attack ranging from 0 to 60 deg. The artificial-viscosity method incorporated with a shock-fitting procedure shows smeared crossflow and wing-shock positions and required 15 percent more CPU per node than the upwind method. Greater flexibility and robustness is demonstrated by the latter on a fixed grid for all cases considered.

Li, Chien-Peng↗