Inviscid hypersonic flow of chemically relaxing air about pointed bodies
Inviscid hypersonic flow analysis for chemically relaxing air about wedges and pointed circular cones using Dorodnitsyn integral method
Hermann, R.↗
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Formulation of image quality prediction criteria for the Viking lander camera
Image quality criteria are defined and mathematically formulated for the prediction computer program which is to be developed for the Viking lander imaging experiment. The general objective of broad-band (black and white) imagery to resolve small spatial details and slopes is formulated as the detectability of a right-circular cone with surface properties of the surrounding terrain. The general objective of narrow-band (color and near-infrared) imagery to observe spectral characteristics if formulated as the minimum detectable albedo variation. The general goal to encompass, but not exceed, the range of the scene radiance distribution within single, commandable, camera dynamic range setting is also considered.
Huck, F. O.↗
Experimental and analytical study of an inlet forebody for an airframe-integrated scramjet concept
Preliminary analytical and experimental inlet forebody investigations have been conducted at Mach numbers of 6.0 and 8.5. The forebody design concept consisted of a sharp-nosed right circular cone followed by elliptical cross sections. This concept resulted in swept isentropic compression which would allow swept cowl leading edges. Measurements were made to define the condition of the inviscid flow field developed by the forebody, including flow profiles in the vicinity of cowl leading-edge stations, and the three-dimensional boundary-layer effects. The investigation verified some of the expected differences between the predicted and the experimental results.
Andrews, E. H., Jr.↗
A higher order panel method for linearized supersonic flow
A higher order panel method using linearly varying source and quadratically varying doublet distributions for computing linearized supersonic flow over arbitrary wings and bodies is described. With tangential mass flux boundary conditions on circular cones at angles of attack, the source panel method gave pressure distributions which are in excellent agreement with exact nonlinear theory. The numerical source panel method was demonstrated to be very stable by randomly panelling an axially symmetric body. For wings with thickness and camber, both the source and doublet panel methods yield excellent agreement with exact linearized solutions. Computational experiments indicate that the use of exact surface paneling in lieu of linearized boundary conditions for wings provides better agreement with higher order theory and with experiment.
Ehlers, F. E.↗
Bi-directional, buried-wire skin-friction gage
A compact, nonobtrusive, bi-directional, skin-friction gage was developed to measure the mean shear stress beneath a three-dimensional boundary layer. The gage works by measuring the heat flux from two orthogonal wires embedded in the surface. Such a gage was constructed and its characteristics were determined for different angles of yaw in a calibration experiment in subsonic flow with a Preston tube used as a standard. Sample gages were then used in a fully three-dimensional turbulent boundary layer on a circular cone at high relative incidence, where there were regimes of favorable and adverse pressure gradients and three-dimensional separation. Both the direction and magnitude of skin friction were then obtained on the cone surface.
Higuchi, H.↗
Unsteady Newton-Busemann flow theory. Part 2: Bodies of revolution
Newtonian flow theory for unsteady flow past oscillating bodies of revolution at very high Mach numbers is completed by adding a centrifugal force correction to the impact pressures. Exact formulas for the unsteady pressure and the stability derivatives are obtained in closed form and are applicable to bodies of revolution that have arbitrary shapes, arbitrary thicknesses, and either sharp or blunt noses. The centrifugal force correction arising from the curved trajectories followed by the fluid particles in unsteady flow cannot be neglected even for the case of a circular cone. With this correction, the present theory is in excellent agreement with experimental results for sharp cones and for cones with small nose bluntness; gives poor agreement with the results of experiments in air for bodies with moderate or large nose bluntness. The pitching motions of slender power-law bodies of revulution are shown to be always dynamically stable according to Newton-Busemann theory.
Hui, W. H.↗
Three-dimensional, time-dependent, MHD model of a solar flare-generated interplanetary shock wave
A three-dimensional time-dependent MHD model of the propagation of an interplanetary shock wave into an ambient three-dimensional heliospheric solar wind is initialized with a peak velocity of 1000 km/s at the center of a right circular cone of 18 deg included angle at 18 solar radii. Differences from a previous 2-1/2 simulation (Wu et al., 1983; Gislason et al., 1984; Dryer et al., 1984) include diminuation of the solar peak velocity and concentration of the peak density at each radius. The IMF magnitude starts with high-latitude peaks, and helical-like IMF rotation is noted due to a large-amplitude nonlinear Alfven wave in the shocked plasma.
Dryer, M.↗
Application of an upwind algorithm to the three-dimensional parabolized Navier-Stokes equations
A new computer code for the solution of the three-dimensional parabolized Navier-Stokes equations has been developed. The code employs a state-of-the-art upwind algorithm to capture strong shock waves. The algorithm is implicit, uses finite volumes, and is second-order accurate in the crossflow directions. The new code is validated through application to laminar hypersonic flows past two simple body shapes: a circular cone of 10 deg half-angle, and a generic all-body hypersonic vehicle. Cone flow solutions were computed at angles of attack of 12, 20, and 24 deg and results are in agreement with experimental data. Results are also presented for the flow past the all-body vehicle at angles of incidence of 0 and 10 deg.
Lawrence, Scott L.↗
Vortex-dominated conical-flow computations using unstructured adaptively-refined meshes
A conical Euler/Navier-Stokes algorithm is presented for the computation of vortex-dominated flows. The flow solver involves a multistage Runge-Kutta time stepping scheme which uses a finite-volume spatial discretization on an unstructured grid made up of triangles. The algorithm also employs an adaptive mesh refinement procedure which enriches the mesh locally to more accurately resolve the vortical flow features. Results are presented for several highly-swept delta wing and circular cone cases at high angles of attack and at supersonic freestream flow conditions. Accurate solutions were obtained more efficiently when adaptive mesh refinement was used in contrast with refining the grid globally. The paper presents descriptions of the conical Euler/Navier-Stokes flow solver and adaptive mesh refinement procedures along with results which demonstrate the capability.
Batina, John T.↗
Thin-layer and full Navier-Stokes, locally-conical and three-dimensional asymmetric solutions
The unsteady, compressible thin-layer and full Navier-Stokes equations are used to numerically simulate steady and unsteady asymmetric, supersonic, locally-conical flows around a 5-deg-semiapex angle circular cone. The main computational scheme used for the present computations is the implicit, upwind, flux-difference splitting, finite-volume scheme. Comparisons of the solutions using the two sets of equations are presented for the flow asymmetry and its control. Computational studies are also presented to investigate the effects of the freestream Reynolds number and the locally-scaled Reynolds number on the flow asymmetry. These studies are carried out using the full Navier-Stokes equations. Three-dimensional, asymmetric flow solutions are also presented for a 5-deg-semiapex angle cone of unit length and a cone-cylinder configuration. The three-dimensional solutions are obtained by using the thin-layer equations and short-duration transient side-slip disturbances along with a very fine grid.
Kandil, O. A.↗
Active control of asymmetric conical flow using spinning and rotatory oscillations
The effectiveness of active control on asymmetric flows around circular cones is investigated computationally using cone spinning and rotatory oscillation around its axis. The investigation uses the time-accurate solution of the unsteady, compressible, full Navier-Stokes equations with the implicit, upwind, flux-difference splitting, finite-volume scheme. The present solutions are obtained under the locally-conical-flow assumption in order to understand the flow physics using very fine grids for reasonable flow resolution at low computational cost. For all the computational solutions, a grid of 241 x 81 x 2 points in the wrap-around, normal and axial directions, respectively, is used. The grid is spinning or oscillating rigidly with the cone according to its motion and the kinematical and dynamical boundary conditions are modified accordingly. The computational applications include the effects of uniform spinning rates and periodic rotatory oscillations at different amplitudes and frequencies on the flow asymmetry.
Kandil, Osama A.↗
Two-dimensional compressible flow in turbomachines with conic flow surfaces
A general method of analysis is developed for two-dimensional, steady, compressible flow in stators or rotors of radial and mixed flow turbomachines with conic flow surfaces (surfaces of right circular cones generated by center line of flow passage in the axial-radial plane). The variables taken into account are: (1) tip speed of the rotor, (2) flow rate, (3) blade shape, (4) variation in passage height with radius, (5) number of blades, and (6) cone angle of the flow surface. Relaxation methods are used to solve the nonlinear differential equation for the stream function. Two numerical examples are presented; one for compressible and the other for incompressible flow in a centrifugal compressor with thin, straight blades. The results of these examples are given by plots of the streamlines, constant velocity-ratio lines, and constant pressure-ratio line.
Stanitz, John D↗
Two-dimensional compressible flow in centrifugal compressors with straight blades
Six numerical examples are presented for steady, two-dimensional, compressible, nonviscous flow in centrifugal compressors with thin straight blades, the center lines of which generate the surface of a right circular cone when rotated about the axis of the compressor. A seventh example is presented for incompressible flow. The solutions were obtained in a region of the compressors, including the impeller tip, that was considered to be unaffected by the diffuser vanes or by the impeller-inlet configuration. Each solution applies to radial and mixed flow compressors with various cone angles but with the same angle between blades on the conic flow surface. The solution also apply to radial and mixed flow turbines with the rotation and the flow direction reversed. The effects of variations in the following parameters were investigated: (1) flow rate, (2) impeller-tip speed, (3) variation of passage height with radius, and (4) angle between blades on conic flow surface. The numerical results are presented in plots of the streamlines and constant Mach number lines. Correlation equations are developed whereby the flow conditions in any impeller with straight blades can be determined (in the region investigated by this analysis) for all operating conditions.
Stanitz, John D↗
Transition Analysis for the HIFiRE-1 Flight Experiment
The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of incidence. This paper reports the initial findings from the ongoing computational analysis pertaining to the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The first mode instabilities were found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear PSE N-factor of approximately 14.
Li, Fei↗
Control of Stationary Cross-Flow Modes in a Mach 3.5 Boundary Layer Using Patterned Passive and Active Roughness
Spanwise-periodic roughness designed to excite selected wavelengths of stationary cross- ow modes was investigated in a 3-D boundary layer at Mach 3.5. The test model was a sharp-tipped 14deg right-circular cone. The model and integrated sensor traversing system were placed in the Mach 3.5 Supersonic Low Disturbance Tunnel (SLDT) equipped with a "quiet design" nozzle at the NASA Langley Research Center. The model was oriented at a 4:2deg angle of attack to produce a mean cross-fl ow velocity component in the boundary layer over the cone. Five removable cone tips have been investigated. One has a smooth surface that is used to document the baseline ("natural") conditions. Two had minute (20 - 40 micron) "dimples" that are equally spaced around the circumference, at a streamwise location that is just upstream of the linear stability neutral growth branch for cross- ow modes. The azimuthal mode numbers of the dimpled tips were selected to either enhance the most amplified wave numbers, or to suppress the growth of the most amplified wave numbers. Two of the cone tips had an array of plasma streamwise vortex generators that were designed to simulate the disturbances produced by the passive patterned roughness. The results indicate that the stationary cross-fl ow modes were highly receptive to the patterned roughness of both passive and active types. The patterned passive roughness that was designed to suppress the growth of the most amplified modes had an azimuthal wavelength that was 66% smaller that that of the most amplified stationary cross- ow mode. This had the effect to increase the transition Reynolds number from 25% to 50% depending on the measurement technique. The application of the research is on turbulent transition control on swept wings of supersonic aircraft. The plasma-based roughness has the advantage over the passive roughness of being able to be adaptable to different conditions that would occur during a flight mission.
Schuele, Chan Yong↗
Quantitative Global Heat Transfer in a Mach-6 Quiet Tunnel
This project developed quantitative methods for obtaining heat transfer from temperature sensitive paint (TSP) measurements in the Mach-6 quiet tunnel at Purdue, which is a Ludwieg tube with a downstream valve, moderately-short flow duration and low levels of heat transfer. Previous difficulties with inferring heat transfer from TSP in the Mach-6 quiet tunnel were traced to (1) the large transient heat transfer that occurs during the unusually long tunnel startup and shutdown, (2) the non-uniform thickness of the insulating coating, (3) inconsistencies and imperfections in the painting process and (4) the low levels of heat transfer observed on slender models at typical stagnation temperatures near 430K. Repeated measurements were conducted on 7 degree-half-angle sharp circular cones at zero angle of attack in order to evaluate the techniques, isolate the problems and identify solutions. An attempt at developing a two-color TSP method is also summarized.
Sullivan, John P.↗
Hypersonic Transition Analysis for HIFiRE Experiments
The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of attack. This paper reports selected findings from the ongoing computational analysis of the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The growth of first mode instabilities is found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear N-factor, based on parabolized stability equations, of approximately 13. Due to the large angles of attack during the re-entry phase, crossflow instability may play a significant role in transition. Computations also indicate the presence of pronounced crossflow separation over a significant portion of the trajectory segment that is relevant to transition analysis.
Li, Fei↗
Stability Analysis for HIFiRE Experiments
The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of attack. This paper reports selected findings from the ongoing computational analysis of the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The growth of first mode instabilities is found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear N-factor, based on parabolized stability equations, of approximately 13.5. Due to the large angles of attack during the re-entry phase, crossflow instability may play a significant role in transition. Computations also indicate the presence of pronounced crossflow separation over a significant portion of the trajectory segment that is relevant to transition analysis. The transition behavior during this re-entry segment of HIFiRE-1 flight shares some common features with the predicted transition front along the elliptic cone shaped HIFiRE-5 flight article, which was designed to provide hypersonic transition data for a fully 3D geometric configuration. To compare and contrast the crossflow dominated transition over the HIFiRE-1 and HIFiRE-5 configurations, this paper also analyzes boundary layer instabilities over a subscale model of the HIFiRE-5 flight configuration that was tested in the Mach 6 quiet tunnel facility at Purdue University.
Li, Fei↗