Swept Cylinder Research Articles

Global Aeroheating Measurements of Shock–Shock Interactions on Swept Cylinder

The effects of the fin leading-edge radius and sweep angle on peak heating rates due to shock–shock interactions were investigated in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. The cylindrical leading-edge models, with radii varied from 0.25 to 0.75 in., represent wings or struts on hypersonic vehicles. A planar oblique shock at 16.7 deg to the flow intersected the fin bow shock, producing a shock–shock interaction that impinged on the fin leading edge. Three fin sweep angles were tested: 0, , and (swept forward). Global temperature data were obtained from the surface of the fused silica fins using phosphor thermography. Metal oil-flow models were used to visualize the streamline patterns for each angle of attack. High-speed zoom-schlieren videos were recorded to show the features of the shock–shock interactions. The temperature data were analyzed using one-dimensional semi-infinite and one- and two-dimensional finite-volume methods. These results were compared to determine the proper heat transfer analysis approach to minimize errors from lateral heat conduction. The dimensional peak heat transfer coefficient augmentation increased with decreasing leading-edge radius. The dimensional peak heat transfer output from the two-dimensional code was about 20% higher than the value from a standard, semi-infinite one-dimensional method.

Investigation of Streamwise and Transverse Instabilities on Swept Cylinders and Implications for Turbine Blading

The starting point for this investigation was the observation of robust streamwise streaks in flow visualization on the suction surfaces of blades in a turbine cascade at subsonic and transonic speeds. The spanwise wavelength of an array of streamwise vortices had been predicted and is here confirmed experimentally. Observations of streaks on unswept turbine blades and on circular cylinders confirmed these earlier predictions, providing a firm basis for referencing the new measurements of vortical behavior. The observations made it clear that the boundary layers are highly three dimensional. In this paper observations of streamwise and transverse instabilities on swept circular cylinders, over a range of inclinations, are presented. The circular cylinder is a canonical case and observations relate the streamwise vorticity of the unswept case to the more aggressive crossflow instability at high sweep angles. Introducing sweep brings consideration of a wide range of instabilities. Prominent is crossflow instability resulting from the inflectional behavior of a three-dimensional boundary layer.

Experimental Investigation of High-Enthalpy Effects on Attachment-Line Boundary-Layer Transition

A series of experiments was performed on high-enthalpy effects on hypersonic boundary-layer transition using nitrogen, air, and carbon dioxideas test gases in the T5 hypervelocity shock tunnel. Previous experiments on a sharp cone showed significant high enthalpy effects. This series concerns the attachment line boundary layer on swept cylinders with sweep angles of 60 and 45 deg. The observed trend of transition Reynolds number with enthalpy, which is found to be similar to that in the cone results, shows strong transition delay at the larger sweep angle for carbon dioxide, whereas no significant effect is observed in nitrogen and air. The acoustic wave absorption rate due to relaxation shows a quite similar trend in enthalpy dependence to transition Reynolds number for the carbon dioxide case. This suggests that the dominant effect in delaying transition in the case is vibrational relaxation. The comparisons between the magnitude of the strongest amplification rate from an inviscid linear stability analysis and the absorption rate due to relaxation show that they are of the same order of magnitude for carbon dioxide and that absorption is not significant for nitrogen or air, which supports the stated understanding of the effect of relaxation on transition.

Stability and Transition on a Swept Cylinder in a Supersonic Flow

Results of experimental investigations of the evolution of natural disturbances and laminar–turbulent transition in a supersonic boundary layer on the attachment line of a circular cylinder with a sweep angle of 68° and a free‐stream Mach number M = 2 are presented. The experimental studies are supplemented by calculations of the mean flow and stability characteristics. Flow regimes in the boundary layer on the attachment line are determined by a hot‐wire technique as functions of the Reynolds number and height of two‐dimensional roughness elements. The results are compared with NASA (Ames) experiments.

Roughness-induced attachment-line transition on a swept cylinder in supersonic flow

Roughness-induced attachment-line transition on a swept cylinder in supersonic flow

Influence of roughness on attachment line boundary layer transition in hypersonic flow

Attachment line boundary layer transition on swept cylinders is studied in a low enthalpy hypersonic wind tunnel at M ∞=7.14. Sweep angles of 60° and 70° are used and transition is detected by means of heat flux measurements. The influence on attachment line transition of single 2D-roughness elements, in the form of tripwires or slots, as well as 3D obstacles is determined and the results are analyzed with respect to Poll’s criterion.

Transition correlation in flow over a swept cylinder

The transition process in boundary-layer flow over a highly swept cylinder is dominated by the crossflow instability. It is of practical interest to have an analytical correlation of transition location in low-speed flow over a swept cylinder with parameters such as freestream Reynolds number and swept angle. We develop such an analytical correlation, based on the extensive expperimental data of Poll.

Boundary-layer transition on swept cylinders at hypersonic speeds

We present boundary-layer transition data on swept cylinders in hypersonic flow that were obtained in the hypersonic Ludwieg-tube wind tunnel of DLR. Experiments were conducted at M∞ = 5.0 and 6.9 for three swept cylinders with and without end plates and sweep angles of 30, 45, and 60 deg, respectively. To determine the state of the boundary layers, the liquid-crystal technique was applied. In addition, the response of attachment-line transition to surface roughness heights due to trip wires for sweep angles of 45 and 60 deg at M∞ = 5.0 was investigated. Freestream Reynolds numbers based on cylinder diameter for attachment-line transition were about 0.2-0.3 x 10 6 for the end plate disturbances and 0.9-1.2 x 10 6 for no contaminations. Trends of attachment-line transition Reynolds numbers with surface roughness heights are roughly similar to subsonic ones with respect to critical roughness heights. The behavior of the attachment-line transition around the critical roughness heights, however, depended strongly on spanwise Mach numbers. With flow visualizations, off-attachment-line transition could be observed, as well as fine streak pattern in laminar flow region upstream of the transition front, confirming streamwise vortices induced by crossflow instability at hypersonic speeds.

Transition in flow over a swept cylinder: Correlation with experimental data

Linear stability analysis, which includes the effects of nonparallelism and body curvature, is used to calculate the N factors of stationary and traveling disturbances for the same experimental conditions measured by Poll for flow over a long swept cylinder. The correlating N factors for traveling disturbances have less scatter than the correlating N factors of stationary disturbances. The mean value of the correlating N factor for traveling disturbances is 14.9 and 7.4 for stationary disturbances. Therefore, for an initial amplitude of a stationary disturbance that is 100 times larger than that of a traveling disturbance, the amplitude of the traveling disturbance is 18 times larger than the amplitude of a stationary disturbance at the location of transition onset. The calculations indicate that transition is due to traveling disturbances; this result is in agreement with the experimental observations.

Unsteady compressible boundary layer flow on the stagnation line of an infinite swept cylinder

Numerical solutions of flow and heat transfer process on the unsteady flow of a compressible viscous fluid with variable gas properties in the vicinity of the stagnation line of an infinite swept cylinder are presented. Results are given for the case where the unsteady temperature field is produced by (i) a sudden change in the wall temperature (enthalpy) as the impulsive motion is started and (ii) a sudden change in the free-stream velocity. Solutions for the simultaneous development of the thermal and momentum boundary layers are obtained by using quasilinearization technique with an implicit finite difference scheme. Attention is given to the transient phenomenon from the initial flow to the final steady-state distribution. Results are presented for the skin friction and heat transfer coefficients as well as for the velocity and enthalpy profiles. The effects of wall enthalpy parameter, sweep parameter, fluid properties and transpiration cooling on the heat transfer and skin friction are considered.

A new hypothesis for transition on the windward face of the Space Shuttle

This paper introduces the hypothesis that roughness induced transition of the attachment-line boundary layers is important for the flow over the windward face of the Space Shuttle. Attachment-line transition is described for the flow over a swept cylinder. The available knowledge is then used to model transition on a slender delta wing at large angles of incidence. This is compared with wind-tunnel data for roughness induced transition on the Space Shuttle. Agreement between the hypothesis and data is good and a simple transition onset criterion is proposed. This is completely general and can, in principle, be applied to vehicles of any shape. Finally, it is shown that the hypothesis provides an explanation for the instantaneous forward flash of the transition front that was observed on the Orbiter Columbia. Nomenclature A = constant b - semispan of delta wing d - diameter of circular trip wire k = characteristic dimension of roughness element L = length of Space Shuttle model M = Mach number Q = resultant velocity (U1 + V2)1/2 R = attachment-line similarity parameter Rk = roughness Reynolds number (Uk - k/v) r = leading-edge radius s = trip (roughness)-detector separation measured along the attachment line T = temperature U, V — velocity components in the x and y directions, respectively

Similar solutions of unsteady, laminar boundary layers on swept cylinders

Fig. 3 Triple point trajectory for a = 90 deg. References 1 Dewey, J. M., Classen, D., and McMillin, D., Photogrammetry of the Shock Front Trajectory on Dipole West Shots 8, 9, 10, and 11, Physics of Fluids, July 1975. Takayama, K., and Sekiguchi, H., Formation and of Spherical Shock Waves in Shock Tube, Rept. of Institute of High Speed Mechanics, Tohoku University, Vol. 43, No. 336, 1981, pp. 89119. Schultz-Grunow, F., Diffuse Reflexion einer Stosswelle, ErnstMach-Institut Bericht No. 7/72, 1972. Bryson, A. E., and Gross, W. F., Diffraction of Spherical Shocks by Cones, Cylinders and Spheres, Journal of Fluid Mechanics, Vol. 10, Pt. 1, 1961, pp. 1-16. Ben-Dor, G., and Glass, I. I., Domains and Boundaries of Nonstationary Oblique Shock Wave Reflections. 1. Diatomic Gas, Journal of Fluid Mechanics, Vol. 92, Pt. 3, 1979, pp. 459-496.

A Correlation of Leading Edge Transition and Heat Transfer on Swept Cylinders in Supersonic Flow

Summary:—It is shown that stagnation line heating rates for laminar, transitional and turbulent boundary layers measured on swept cylinders at supersonic speeds, can be correlated in terms of a Reynolds number based on the spanwise momentum thickness at the stagnation line.A critical Reynolds number for transition along the stagnation line is determined and experimental results show evidence of transition up to M= 10.